Current Space Shuttle System Shortfalls Assessment · 10/18/2005 · Current Space Shuttle System “Shortfalls Assessment” 4 9/28/2010 the space transportation system LCC must

Report of

Current Space Shuttle System

Shortfalls Assessment

Produced by

THE SPACE PROPULSION SYNERGY TEAM’S (SPST)

FUNCTIONAL REQUIREMENTS SUB-TEAM

For the:

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION

HEADQUARTERS OFFICE

Director Strategic Investments

Revised: October 18, 2005

Current Space Shuttle System “Shortfalls Assessment”

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Table of Contents

Executive Summary Page 3

1.0 Introduction Page 5

2.0 The SPST Approach Page 6

3.0 Background (SPST Supporting Analysis and Studies) Page 6

3.1 Space Shuttle Page 6

3.1.1 Space Shuttle Level 1 Program Requirements Documents Page 6

3.1.2 Space Shuttle History Overview Page 6

3.1.3 Space Shuttle Fluids Overview Page 7

3.1.4 Hardware Dependability Influence on Propellant Cost/LCC

Goals

Page 7

4.0 Space Shuttle Shortfalls and Future STS Needs Overview Page 7

4.1 Space Shuttle Shortfalls Page 10

4.2 Future STS Needs Page 23

4.3 Fluid & Propulsion System Technical Needs (Flight & Ground Sys) Page 25

4.4 Balancing Safety, Reliability, and Maintainability Requirements Page 26

4.5 A Generic Functional Systems Breakdown Structure (SBS) for Space

Transportation Architectures

Page 29

5.0 Summary Page 29

6.0 Conclusions Page 31

7.0 Recommendations Page 31

Appendix I – Space Shuttle Level I and II Program Requirements

Documentation excerpts

Page 32

Appendix II - Space Shuttle Launch Information Data Base Page 45

Appendix III - Space Shuttle Fluids Overview Matrix Page 51

Appendix IV - Hardware Dependability Influence on Propellant Cost/LCC

Goals

Page 52

Appendix V - Shuttle Reference Case 1: safety driven functional

requirements

- Number of safety driven functional requirements to maintain

safe control of systems during flight and ground operations

Page 58

Appendix VI - Shuttle Reference Case 2 for safety driven limited access

control

- Number of safety driven limited access control operations

Page 60

Appendix VII - Space Shuttle Shortfalls Assessment Results Page 61

Appendix VIII – Summary of Annual Shuttle Payload Delivered or

Retrieved from Space

Page 81

Appendix IX - Glossary Page 82


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Executive Summary

Recommendations. It is clear from the ―Shortfall Assessment‖ performed by the national space

propulsion synergy team (SPST) and summarized in this document, that past and current Shuttle

Program efforts to control life cycle costs (LCC) have been inadequate and ineffective.

Therefore, the (SPST) is recommending NASA consider adopting the proven methods of

controlling weight and performance and applying them to controlling cost (LCC). It is also

recommended that the NASA consider using the SPST technique for balancing Safety,

Reliability, and Maintainability requirements to provide controls on recurring maintenance

burden to provide operational effectiveness and LCC. In addition, the future programs need a

complete functional systems breakdown structure (Functional SBS) of all activities to provide

the total visibility of the entire task. This functional SBS will allow the systems engineering

activities to totally integrate each discipline to the maximum extent possible and optimize at the

total system level. Optimizing at the individual sub-system level creates a very large support

infrastructure (very large labor workforce) associated with resultant high LCC.

The SPST recommends that the Program Analysis and Evaluation Office endorse these

SPST recommendations and implement requirements for all new NASA programs to

require these LCC and operational controls. Further the SPST recommends that these

new approaches be implemented immediately within the current planning of the Space

Exploration Program Missions.

We emphasize these recommendations, because, the Space Exploration Program must not

only be ―affordable‖ but ―sustainable‖. This requires close control of life cycle costs

within established budgets. In addition, this effort will help re-gain NASA’s cost

creditability with the country’s leaders.

Objective of this Report. The ultimate objective of this report is to assure that the planning and

implementation of the transportation systems required by the Space Exploration Program and

any future systems takes maximum advantage of the ―lessons learned‖ from the major space

programs of the past decades. The focus of the report is on what has been learned about the

assessment and improving control of Life Cycle Costs (LCC) from these major space programs.

The major ―lesson learned‖ from these studies is that much improved, innovative processes must

be developed and rigorously applied to effectively control life cycle cost. It is also learned that

NASA must balance the Safety, Reliability, and Maintainability requirements to provide controls

on recurring maintenance burden to provide operational effectiveness and LCC.

The only major objective that was controlled in these past programs by a structured Engineering

Management process was performance closure by managing flight systems weight. Objectives

were set for Life Cycle Cost (LCC) for the Shuttle, but no Engineering Management processes

were exercised to provide control (only the DDT&E cost was tracked but not controlled).

For example, the Saturn/Apollo lunar exploration program was terminated early because the

recurring transportation cost was not sustainable while supporting the exploration efforts. The

reusable Shuttle transportation system was developed to replace the Saturn launch vehicle in an

effort to greatly reduce the recurring cost of transportation. Therefore, the lesson learned was that


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the space transportation system LCC must be controlled to provide a sustainable space

exploration program. The major part of the space transportation LCC is the recurring or

operational phase cost.

To accomplish this critical objective, this report provides the results of the current space

transportation system shortfalls assessment and supporting analyses that have been conducted by

the Space Propulsion Synergy Team (SPST). These studies directly address the ―lessons learned‖

from previous transportation systems and the way NASA does business, as well as

recommended/suggested solutions for improvement. The SPST has a proposed option to control

LCC by controlling the major operational technical functions that greatly influence LCC through

the use of requirements and Engineering Management control processes.

Lessons Learned from SPST Studies. A major ―lesson learned‖ from these activities is the

importance of first clearly defining, flowing down, and controlling the ―systems requirements‖

and maintaining control throughout the DDT&E Program. The SPST has emphasized the need to

clearly define the ―requirements‖ up front: that is the ―what’s‖ required of the desired space

transportation system. These requirements must cover all major objectives, not only

―performance‖, as was the case in the past, but also in terms of the ―functional (operational)

requirements‖ required in the system to achieve sustainable Life Cycle Cost, safety and the

country’s support. To sum up this lesson learned, we must change the way we do business to

avoid ―doing what we always do and achieving what we always got‖. Therefore, we must change

our Engineering Management processes to include a structured process to control those major

operational functions that are major cost influences to provide the LCC controls required for a

sustainable Space Exploration Program – total program optimization.

Insight gained from performing the shortfalls assessment stresses the need to perform

optimization at the total systems level and not at the sub-system level (stove-piping). The SPST

has developed a new approach for formulating ―requirements‖ that will provide full

accountability of all functions required to perform the planned space missions. The approach as

briefly described in this report was to develop a top-level functional systems breakdown

structure, (Functional SBS) with modular sub sets, that may be utilized as a basis for defining the

desired ―functional requirements‖ in any space system. This process is intended to serve as a

guide in development of the work breakdown structure (WBS), provide visibility of those

technologies that need to be developed to cover a required function, and help identify the

personnel skills required to develop and operate the space transportation system for this very

large and challenging National effort. This Functional SBS covers all transportation elements on

earth, the moon and mars including any orbiting operational space nodes if deemed necessary.

This SPST report is available under separate cover.

Another study performed by the SPST was a ―bottom-up‖ analysis which addressed the question

of why past programs weren’t achieving the desired functional criteria: ―what has impeded or

prevented the application of good systems engineering and management’s successful

implementation of the approaches/processes addressed in this report?‖ Results are very

stimulating and deserving of more in depth attention. For example, it was found that there are

several reasons for the impediments: lack of overall integration (stove-piping or optimizing at the

single function level), inappropriate starting technology level, the lack of sufficient Engineering


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Management processes, and that many of the systems engineering requirements (needs), were

―boring‖ not stimulating (not sexy). This indicates that major improvements in discipline must be

rigorously imposed on the system engineering and design processes by the program managers

and the chief systems engineer. This effort is not included in this report, but is available on the

SPST web site at the NASA MSFC Virtual Research Center at Huntsville (VRC).

1.0 Introduction

In the spring of 2002, Dan Dumbacher, the Second Generation Reusable Launch Vehicle

(2GRLV) Deputy Program Manager, requested that the SPST review and critique the

2GRLV Program Level I Requirements documents. He specifically was interested in a

critique of these requirements regarding their capability of assuring the design,

development and eventually the operation of a Space Transportation service that would

meet the 2 GRLV Program goals/objectives. It was stated in the 2GRLV Requirements

document that the 2GRLV Level I architecture requirements are derived from an

understanding of NASA, commercial, and projected DOD mission needs, and from

identified areas for improvement gleaned from experience with the existing fleet of

launch vehicles. The 2GRLV program personnel, working to a short schedule chose to

by-pass the assessment of existing launch vehicles for shortfalls and to identify the needs

for improvement. After some time, the SPST Requirements Sub-team agreeing this was

an important step in the identifying of requirements for a next generation space launch

system, undertook this task. Being a mostly volunteer task, it has been aborted several

times for higher priority work and is just now being formally completed.

Civil and military applications of Space Transportation have been pursued for 50 years and there

have been and there is now an even greater need for safe dependable affordable and sustainable

space transportation systems. Fully expendable and partially reusable space transportation

systems have been developed and put in operation. Access to space is technically achievable, but

presently very expensive and will remain so until there is a breakthrough in the way we do

business. The approach to providing the propulsion systems functions has a major influence in

achieving the affordable/sustainable objectives and again will require a breakthrough in the way

NASA has been doing systems engineering and management.

A critical need for improved communications between the user and the developer led to NASA’s

Code R and Code M chartering the Space Propulsion Synergy Team in 1991. The SPST’s first

task was to use its member’s diversified expertise toward developing new ―Engineering

Management Decision Making Tools‖: specifically developing innovative engineering processes

in the architectural design, development, and operation of space transportation systems to satisfy

the challenging requirements of both the transportation operators and the payload customers.

The SPST established a dialogue between the personnel involved in all phases of the technology,

design, development, and operation of a space transportation system.

This report describes the development of the process used to document the current space

transportation system shortfalls providing a clear visibility for the need to improve on both the

technical and management processes required to achieve the objectives desired of a next

generation space transportation system.


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The basic approach by SPST to the task of providing defining and documenting the current

system shortfalls is to compare the achievements to the original requirements, objectives, and

goals of the system being provided. This includes technology, advanced development, DDT&E,

Manufacture, Operational, and Recycle/Disposal phases of the current program.

The SPST proposes to address the global problem of ignoring past program lessons learned,

technical issues that get passed from one program to the next, and not taking these shortfalls into

consideration in developing the architectural concept along with its requirements formulation by

providing this formal report to the responsible decision maker at the start of the major

undertaking of providing a replacement space transportation system.

2.0 The SPST Approach

The objective of the study was to define the major program shortfalls for the Space Shuttle

Transportation System both programmatic and technical. The approach was to define and

document the major program requirements/objectives/goals at the first and second levels. The

next step was to define and document the accomplishments to each of the

requirements/objectives/goals documented. The shortfalls were then documented along with

need statements for improvement to the next generation Space Transportation System. This data

was analyzed to determine the cause for the shortfall so that the appropriate corrective action

could be identified. This effort could be stated as the lessons learned from the Space Shuttle

program.

3.0 Background (SPST Supporting Analysis and Studies)

3.1 Space Shuttle

Although the Space Shuttle is a highly successful program, the first of its kind, and has

produced cutting edge technology, it has many requirement, objective, and goals

shortfalls. By looking at its history, the SPST determined what were its major objectives

and goals and what was actually achieved. The need to focus on specific areas was made

visible and a number of "lessons learned" were derived and are presented in this report.

3.1.1 Space Shuttle Level 1 Program Requirements Document

Space Shuttle Program Level 1 documentation excerpts of interest can be found in

Appendix I. This copy is the last revision maintained in original context before revising

to show conformance to achievement. This document will also serve as the primary

Requirements, Objectives, and Goals reference for this shortfalls analysis.

3.1.2 Space Shuttle History Overview

A historical perspective of many of the shuttle program operations parameters were

established to provide visibility and allow insight to shortfalls in the program. This

shuttle operations perspective matrix can be found in Appendix II.


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3.1.3 Space Shuttle Fluids Overview

Insight into one of the operational characteristics (number of different fluids required)

was gained by building a matrix of all the fluids used by the Shuttle launch vehicle and

providing visibility to their sub-systems use and its cleanliness requirements. This

information shows there are 102 total unique sub-systems that require fluid servicing for

each launch and several using the same commodity (17 separate He servicing locations).

This assessment indicates a total lack of integration of flight functions to require the

minimum ground infrastructure, flight hardware logistics support, and minimum

sustaining engineering. Composite matrix can be found in Appendix III.

3.1.4 Hardware Dependability Influence on Propellant Cost/LCC Goals

This analysis provides insight that the depot maintenance function drives the propellant

cost for Shuttle to exceed the flight operations propellant costs. It also indicates that the

design life and reliability relationship was not sufficient to meet the expected

dependability goals; therefore, driving up the LCC or cost per flight expectations beyond

its goal. The insight provided here is that the safety, reliability and maintainability

requirements must be balanced to provide controls of recurring cost. Shuttle experience

indicated that the recurring cost was left to chance because the maintenance burden

(depot cycle) was not controlled by a flow down of requirements. This complete analysis

can be found in Appendix IV.

4.0 Space Shuttle Shortfalls and Future STS Needs Overview

NASA Mission and Space Transportation’s Role

NASA’s Vision:

- To improve life here

- To extend life to there

- To find life beyond

NASA’s Mission:

- To understand and protect our home planet

- To explore the Universe and search for life

- To inspire the next generation of explorers

…..As only NASA can

Today, NASA utilizes both the Space Shuttle and expendable launch vehicle

transportation systems in accomplishing its mission. Assessments of future needs

indicate capabilities of current space transportation systems are not adequate and

additional capabilities will be required to accomplish NASA’s mission in the near,

mid and far term. Current space activities are constrained by a lack of operational

flexibility and responsiveness, high cost, limited in-space maneuver capability,


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concerns for safety and reliability, and significant constraints on payload mass and

volume.

Current RLV Space Transportation Systems Shortfalls Identification

In order to identify the requirements needed to develop the systems and the supporting

technology for future space transportation systems (near, mid, and far), first the

shortfalls must be identified with the current systems. The shortfalls must be identified

and analyzed to determine the entire root causes responsible. These shortfalls’ root

causes must be corrected by developing the needs statements required to identify the

global requirements for future space transportation systems. These top level

requirements must be followed by the development of the next level requirements that are

required to be in place to influence the architectural concepts that will satisfy and correct

all the shortfalls of the current systems.

In identifying the shortfalls of current systems all of the desired functional attributes of

the space transportation system should be addressed. These functional attributes desired

of these systems are listed below.

Functional Attributes Desired of the Current System:

1. Affordable / Low Life Cycle Cost 2. Very Dependable System

1.1 Low initial acquisition cost 2.1 Highly Reliable Hardware

1.2 Low Recurring cost 2.2 High Intact Vehicle Recovery Rate

1.3 Low Vehicle/Sys Replacement Cost 2.3 High Mission Success Rate

1.4 Low Sensitivity to Flt Growth Cost 2.4 Very Robust

1.5 Low Direct and Indirect Labor Cost 2.5 Always Operates on Command

1.6 Min Payload Cost Impact on Launch Sys 2.6 High Design Certainty

1.7 Must Close Commercial Bus Case

3. Highly Responsive to Space Customer’s Needs 4. Safety of Hardware and Personnel Very High

3.1 Very Flexible to Destination Locations &

Manifesting (Capable of surge and launch on

demand)

4.1 Flight & Ground Hardware safety derived from

highly reliable & maintainable parts/systems

3.2 Responsive to Desired Capacity Range 4.2 Public & Personnel Safety derived from Very

Reliable/Maintainable Flight & Ground Systems

3.3 Minimum and Ease of Process Verification 4.3 Environmental Safety/Compatibility with

Space, Atmosphere, and Ground

3.4 Designed in Automatic Functional Health

Verification 5. Low Acquisition Risk Compatible with

Investor’s Incentive

3.5 Designed in and Automatic System Corrective

Action

5.1 Technology Maturity Very High for ALL

Systems

3.6 Vehicle/ Payload/Ground Integration Ease

5.2 Acquisition Cost Well Understood and

Compatible with Commercial Business Case Closure

3.7 Maintainable (minimum and very accessible

maintenance) 5.3 Technology Options Available and Mature at

Acquisition Start (No open R&D required existing)

3.8 Labor Skills Required to Operate Low 5.4 Acquisition Schedule Compatible with

Commercial Business Case

3.9 Easily Supportable (minimum support

infrastructure required)


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The ―US Space Shuttle Shortfalls‖ provided the rationale for Space Transportation Systems Life

Cycle Cost Assessment and Improving Control to be completed early so the Life Cycle Systems

Engineering and Integration (SE&I) Life Cycle Cost (LCC) disciplines can be applied in the

initial definition of the space exploration program establishing LCC as the major and overriding

metric for all hardware and software implementation programs.

The Space Shuttle is not adequate to accomplish the spectrum of NASA-missions goals

(not sustainable) because current space activities are constrained by the following:

Operational flexibility and responsiveness – flight rate has not achieved concept

goals.

Operated by RDT&E personnel—the developer (instead of ―commercial-style‖

operations personnel) with resultant high operations cost – there is no reward

incentive, or system, to support ―order-of-magnitude‖ cost cutting.

Limited in-space maneuver capability – science and logistic mission scopes are not

all-inclusive of agency vision.

Concern for safety and reliability is constrained to the system architecture – what

you see is what you get.

Significant constraints on payload mass and volume – greater ―Operability‖ (flight

rate) is needed to reduce historical LCC ($/PL lb to orbit/year) and provide much

larger annual mass-throw capability; i.e., the learning curve.

The SPST Functional Requirements Sub-team prepared a table which shows the current

capabilities of the Space Shuttle and the Critical Shortfalls relative to the initial Space Shuttle

requirements. This table can be found in Appendix I. The attribute ID column in the table is a

reference back to the previous table of attributes by function and their number, e.g., A 1.2 would

refer to the Affordable attribute and item ―1.2 Low Recurring Cost‖.


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4.1 Space Shuttle Shortfalls

SUMMARY OVERVIEW

Current RLV Space Transportation Systems Shortfalls Assessment

Space Shuttle Critical Shortfalls Relative to Requirements

Attribute

I D.

Program Objectives and

Desired Attributes needed

to Correct Shortfalls of

Current Systems

Target Value Actual Achieved

Critical Shortfalls

Relative to Requirements

D 2.1 1. Program STS Design Life

Propulsion Main Engine

Life

10 years or

100 flts/veh.

55 starts/Depot

cycle.

20+ years, but only

30 flts/veh. Max., but still

counting

After 20+ years ops,

SSME depot cycle 20,

LO2 turbo-pumps 10 &

fuel turbo-pumps 3 flts

The Space Shuttle was intended to fly 10 flights

per year each without extensive maintenance and

recertification between flights (160 hour turnaround).

Design complexity and hardware dependability only

permits less than 3 flights per year. Avg. 100

components replacement plus ~ 400 expendable

or limited life parts. The SSME initial design life

was 55 flights before entering depot cycle, but

limited life/dependable hardware has required

extensive labor, time, and engine depot support, e.g.,

resulting in high cost per flight. Application must be

well understood so that the reliability requirements

flow-down supports the design life after balancing

the requirement with safety and maintainability. Also

the reliability requirement must be demonstrated by

testing and improved until the requirement is met.


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Attribute

I D.




Current Systems


Critical Shortfalls


A 1.2

S 4.3

2. Recurring cost:

All processes and

operations must be

compatible with

environmental

regulations and laws

$100.00/lb to

orbit

No Requirement

documented?

Today’s

regulations were

not established

during the Shuttle

concept and

DDT&E phases

and it was

assumed NASA

would abide by

the country’s

laws.

~ $10,000.00/lb to orbit

Actual Shuttle recurring

cost over the total 21

operating years =

~ $57.876 Billion.

Stringent Environmental

/OSHA requirements have

been imposed since Shuttle

ATP

The initial design recurring costs were

$100.00/lb to LEO which was based on achieving

an allocated 40 launches per year using 4 orbiters

flying 65,000 lbs per flight. Most flights do not fly at

maximum capacity and the 10 flights per year per

each orbiter was not achievable because of

complexity (optimizing at the sub-system level

―stove-piping‖ and not at the overall STS systems

level) and poor dependability of total system.

STS Program didn’t include cost allowances for

changes in the environmental laws.


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Attribute

I D.




Current Systems


Critical Shortfalls


A 1.1/

P 5.2,

P 5.3

& 5.4

A 1.7

3. Non-recurring

cost:(DDT&E and

Acquisition)

LCC must be well

defined and understood

by analysis without any

allocations/assumptions

so that the business case

closes

All technologies must be

matured at the TRL-6

level or above and

options must be

available as backup

where risk is moderate or

above prior to the start of

acquisition.

$5.0 Billion

The targeted NRC

was $5 billion

and the Recurring

target was $6.5

million/flight = to

$2.6 billion in 10

yrs. Or a total

LCC = to $7.6

billion.

Assumptions

were 65,000 lbs

to orbit each

flight, 10

flights/orbiter or

40 flights/yr at

$100/lb to orbit.

DDT&E schedule

and cost risks

were not

considered a

necessity as we

were still working

to the Apollo

paradigm.

Shuttle NRC (DDT&E) =

$15 billion and the RCC

average is ~$2.756

billion/yr. Therefore, the

intended 10 year program

LCC would have = $42.56

billion. But the Actual

Shuttle recurring cost over

the total 21operating years

= ~ $57.876 B. Or a total

LCC = $72.876 B.

These actual cost do not

include any R&T cost prior

to the STS ATP (1-5-72),

e.g., SSME, TPS, etc.

Five Major System’s

Technologies less than

TRL-6 level at ATP: High

Pressure LO2/LH2 Staged

Combustion Rocket

Engine, Vehicle TPS,

Large Solid Rocket Motor

Nozzle Flex Seal TVC

system, Ice/frostless

cryogenic tanks, & 100%

digital flight/ground

control systems

The initial design non-recurring cost

estimate were $5.0B based on an allocated DDT&E

schedule. Due to non-mature major technologies

(HP LH2/LO2 staged combustion rocket engine,

re-entry TPS, Solid rocket flex nozzle seal, Ice/frost

less cryogenic tanks, and 100% digital flight/ground

control systems), schedule was overrun 2 years

because much unplanned technology maturation was

required. Started the development with high risk

schedule for technology maturation without providing

a margin in cost or schedule to account for this high

risk approach. There were no requirements or policy

documented towards the use of mature or non-mature

technologies.

A very large shortfall exist in the LCC projections

because they were based on allocations that never

came into fruition, e.g., 65,000lbs to orbit each flight

and 40 launches per year using 4 orbiters. Also the

DDT&E cost projection had a large shortfall

because of the immature technologies causing an

extended schedule for this activity. Allocations of

the operational functions could not be met because

there was no engineering management processes in

place to provide the necessary control required.


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Attribute

I D.




Current Systems


Critical Shortfalls


R 3.2

R 3.1

&3.9

D 2.6

4. Each vehicle flight rate:

Fleet flight rate:

Vehicle turnaround

time:

System performance to

LEO

Materials, fluids, and

design properties and

limitations well

understood through

failure with narrow

tolerances

10 flights/yr.

40 flights/yr. at

ETR

160 hours

65,000 lbs at ETR

@ 28.8 degrees

and 100 nmi

Considered as

over constraining

and would have

driven up the

DDT&E cost

considerably.

2.5-3 flights/yr.

10 flights/yr.

1296 hours Min.

55,000 lbs at ETR @ 28.8

degrees and 100 nmi

Because the limits were not

known, operational

controls provided margins

to avoid unplanned events.

Performance carried an

extra margin to allow for

these uncertainties.

The initial design allocation for turnaround was

160 hours landing to re-launch. The initial design

flight rate was 10 flights per year for each orbiter,

but because design requires the functional integrity

to be broken each flight to perform the turnaround,

~ 400 expendable or limited life parts to be replaced,

and ~ 100 failed components to be replace during

the turnaround operation, and extensive servicing

(too many different fluids & too many interfaces

along with the extensive support infrastructure)

required the achievable flight rate is just above 2

per year. Because the integrity of systems are

compromised to provide for parts change-out and the

support turnaround operations, the STS must be

re-certified for each flight. The shortfall in payload

mass capacity was a result of lack of sufficient

margins in performance of each variable, e.g.,

orbiter over weight, SSME Isp low, and the drive to

keep the ET production cost low. Example of cost

to remove ET weight was and additional

$20,000,000. /unit for a 6,000 pound reduction.

Program objectives were compromised because of

the added limitations do to uncertainties.


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Attribute

I D.




Current Systems


Critical Shortfalls


S 4.2

R 3.1

&3.9

D 2.5

S 4.1

&4.2

5. Space flight rescue

time:

Launch Availability:

Vehicle/System terminal

countdown:

Launch on time (No

launch scrubs)

Flight abort during

ascent:

No Loss of Crew

Flight abort from orbit:

No Loss of Crew

24 hour notice to

launch from

standby status

(VAB/T-0) VAB

rollout including

payload change-

out at the pad,

MPS propellant

loading, crew

ingress, final

close-out checks

and terminal

count.

2 hours

No Requirement

documented

except 24 hr.

notice to launch

for space rescue

& military needs.

Designed for

RTLS/ATO/AOA

KSC prime,

EAFB secondary,

& several

contingencies

Not capable, but now being

considered again since

Columbia event.

14 Work Days at the Pad is

best case before STS-51L

and 19 Work Days at the

Pad has been demonstrated

after the STS-51L event.

8 hours plus

65 of the 113 missions

launched the day scheduled

(57.5%). Of the 48 launch

scrubs, 13 were weather

related (27%) However,

some missions were

scrubbed more than

once/mission.

Did not demonstrate

RTLS or TAL’s and

ATO was required only

once, but did not result in

an aborted operation. 3

landing sites used: (KSC,

EAFB, & White Sands)

The requirement for the 2 hour terminal countdown

was deferred because of the added DDT&E cost to

provide the automation for crew egress and MPS

Lox transfer capacity needed and the lack of meeting

the fleet flight rate.

There was no requirement against reliability to

accomplish either the launch on time or meet the 24

hour notice to launch for a space rescue. Not

considered as a need to provide any control and was

considered as over constraining.

Requirements flow-down were not developed,

implemented and controlled to provide this

capability. Lack of major system integration

resulted into too many flight/ground service

interfaces, controlled access conditions and

extensive time consuming operations.

Abort during ascent operations required the SRB’s

to burn to completion and failure occurred with the

SRB resulting in the loss of the orbiter (099) and its

crew. Therefore, abort during ascent did not cover

all critical failure modes.

No abort was provided during the descent phase and

an orbiter (102) and its crew were lost during re-entry.

The STS vehicle reliability wasn’t sufficient to

support the abort modes required.


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Attribute

I D.




Current Systems


Critical Shortfalls


D 2.1

D 2.2

D 2.3

D 2.4

&

S 4.1

&

R 3.1

6. STS

Dependability/Safety

Loss of Vehicle

Flight system program

reliability:

Mission reliability:

Flight environment:

Launch & Landing

All flight vehicle

subsystems (except

primary

structure, thermal

protection system, and

pressure

vessels) shall be

established on an

individual

subsystems basis,

but shall not be

less than fail-safe.

Safety, reliability,

and maintainability

were controlled

separately by NHB

5300.4 (lD-1),

August 1974, or

0.98 for 100

missions of each

orbiter or 500

missions total for

fleet of 5 orbiters.

Requirement for

95 percentile

natural

environment

expected at

operational

locations

Program has lost 14 flight

crew and 2 ground

members and two orbiters,

e.g., 0.964 for the Orbiter

and 0..962 for the SRB and

a mission

Reliability of ~0.96

7 mile visibility & no rain

Program did not consider cost impact of vehicle loss

accompanied with down time for the investigation

and corrective action required for re-flight.

Importance of loss of vehicle and the resultant

impact on the program wasn’t considered with

proper risk reduction actions. Target metric value

for reliability was deficient in determining its overall

judgment in importance. No requirements were

established for loss of flight or ground crew members

and the impact of insufficient component reliability

was not considered and understood. Target metric

value was also deficient in determining its overall

impact on the maintainability burden (plus large

depot maintenance and supply chain support)

resulting in reduced flight rate and increased cost per

flight. Because the recurring cost per flight was not

controlled, the mission reliability importance was not

understood. Orbiter TPS cannot function in design

environment without damage. TPS needs to be more

robust to be in compliance with requirements and to

avoid launch and landing scrub/delayed operations.

This lack of robustness attributed to the loss of an

orbiter and 7 crew members.


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9/28/2010

Attribute

I D.




Current Systems


Critical Shortfalls


R 3.7

D 2.1

7. STS

Dependability/Safety/

Maintainability:

Component replacement

time or MTTR:

Shuttle orbiter

was designed for

100 flights or 10

years without

planned

maintenance. No

other Direct

Requirement

other than (160

hours turnaround)

Except the Shuttle

SSME 55

starts/Depot

cycle. SRB was to

be recovered and

refurbished every

flight.

Replace Avg. of 100

components/flight

unplanned & best case

orbiter turnaround is ~ 960

hrs. There are many limited

life components on the

Shuttle orbiter, e.g., ~ 200

expendable ordinance

items and ~ 200 other

limited life items to track

& replace.

Also after 20+ years of

ops, SSME depot cycle is

every 20 flights, with the

LO2 turbo-pumps after

every 10 & fuel turbo-

pumps after every 3 flights

Example of SSME MTTR

Controller replacement

during scrub-turnaround:

Up to 5 days or 80 Hrs.

Only requirement was the 160 hour turnaround and

maintainability design efforts were dropped early in

the DDT&E phase because of cost overruns and

schedule concerns. Critical component redundancy

was implemented with component reliability levels

that ignored the resultant maintainability burden.

This lack of controlled maintainability requirements

(accessibility, intrusive nature of most of the

hardware and no automated functional verification)

has contributed to the large resultant cost per flight

and the low flight rate. Controlled maintainability

requirements properly balanced with safety and

reliability using existing methodologies is

major shortfall in the STS program and has contributed

to the large resultant cost per flight and the low flight

rate.


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9/28/2010

Attribute

I D.




Current Systems


Critical Shortfalls


R 3.9/

A 1.5

8. Total number of

assembly functions

required at the launch

site between flights

The shuttle initial

design requirement

provided an

allocation of 34

hours of the 160

turnaround for the

space vehicle

assembly. Also a

40 flights/yr. fleet

flight rate with 4

vehicle fleet size.

The two SRB stages are

completely assembled from

scratch at the launch site

for each launch on the

MLP, a new ET is received

and integrated into the

SRB/MLP stack, with the

Orbiter being integrated as

the final step of building

the flight vehicle. The

Orbiter requires re-

configuration for each

unique payload structural

attachment as well as

providing unique airborne

support equipment to

service the payload after

installation into the Space

Vehicle.

The large SRB vehicle element concept does not

lend to the objectives of an RLV that achieves a 40

launch per year flight rate as it must be built-up at

the launch site and the recovery operations are more

like salvage and reconstruction operations. Design

concept choice was inappropriate for the objective of

the space transportation system.

R 3.9/

A 1.5

D 2.1

9. Total number of

expendable

items/components

included in the reusable

system design

Objective was to

provide an RLV

with a 160 hours

turnaround

capability that

could fly 40 time

a year with a fleet

of 4 orbiters.

~ 200 ordinance items

replaced every flight and ~

200 other one-flight limited

life items on the orbiter

plus the expendable ET

and much expendable

hardware on the SRB’s.

Designers and program managers loss sight of the

objective to build an RLV because of an overriding

focus on meeting the performance requirements in

the absence of any other structured engineering

management processes used to control such thing as

life cycle cost or any of the other program level one

objectives.


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Attribute

I D.




Current Systems


Critical Shortfalls


R 3.9/

A 1.5/

R 3.6

10. Total number of sub-

systems requiring

servicing with dedicated

ground systems, e.g.,

number of different

Fluids, number of

different Electrical

supplies, etc.

Easily Supportable

(minimum support


and must be compatible

with all LCC

requirements by analysis

without any assumptions

Objective was to

provide an RLV

with a 160 hours

turnaround

capability that

could fly 40 time

a year with a fleet

of 4 orbiters and

with a 24 hr.

notice to launch

capability to

accommodate

rescue.

Shuttle System requires the

tracking and managing of ~

54 different fluids and ~ 30

unique fluids are serviced

every flight. Many of these

fluids are common from

one discipline to another,

which require separate

umbilicals, as they do not

share storage on the

vehicle. The Shuttle has ten

(10) major sub-system

disciplines that require

fluid servicing between

flights with several unique

support systems that also

require servicing every

flight. The total of 102

dedicated sub-systems

requires servicing for each

flight. Seventeen (17)

dedicated electrical power

supplies that required

support and service each

flight.

STS cost analysis was shallow and was based on

allocations with no follow-up in establishing

requirements flow-down to assure compliance in

meeting these objectives. LCC analysis must be

realistic and based on the functional requirements

along with assurance the objectives can be met.

Not considered was a need to provide any control

that would have been considered over constraining.

Designers and program managers loss sight of the

objective to build an RLV because of an overriding

focus on meeting the performance requirements in

the absence of any other structured engineering

management processes used to control such thing as

life cycle cost or any of the other program level one

objectives. Also the STS was optimized at the sub-

system level ( stove-pipe approach) and not at the

overall integrated level. Electrical functions are

custom managed on the ground and uniquely

provided through separate umbilicals instead of

simplifying the flight to ground interface functions

by providing the electrical management on the

vehicle. Major shortfall is the need for structured

engineering management process (like the one used

by Shuttle to control weight/performance) to provide

controls that would drive overall system integration.


19

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Attribute

I D.




Current Systems


Critical Shortfalls


R 3.9/

A 1.5/

R 3.6

10. Total number of sub-

systems requiring

servicing with dedicated

ground systems, e.g.,

number of different

Fluids, number of

different Electrical

supplies, etc.

Easily Supportable

(minimum support


and must be compatible

with all LCC


without any assumptions

Con’t

Data bus and

communication systems as

well as unique

instrumentation have not

been accounted for in this

assessment.

Orbiter element alone has

402 functional interfaces:

Propulsion discipline has

236 of which an SSME has

25/engine documented in

the formal structured flight

to ground interface (ICD)

system for the single

ground turnaround facility

station (ICD-2-1A002).

Note: The orbiter element

has ten (10) more facility

station ICD’s at the launch

site.


20

9/28/2010

Attribute

I D.




Current Systems


Critical Shortfalls


R 3.6

R 3.9

A 1.5

11. Degree of custom build

required to support each

mission

Total number of manual

functions required to

determine and control

critical flight functions,

e.g., CG, fluid residuals

content & purity,

functionality of primary

and backup system

hardware

Vehicle, payload, and

ground systems

integration functions

must be compatible with

all LCC requirements by

analysis without any

assumptions.

The shuttle initial

design requirement

provided an

allocation of 96

hours of the 160

turnaround for the

orbiter turnaround

including the

payload installation and

verification. Also a


flight rate with 4

vehicle fleet size.

Each different payload

requires the Orbiter to be

custom built to support the

structural load and any

servicing requires special

airborne support equipment

to be installed and verified

along with optimizing the

mass impact on the

payload for these services.

Also flight software must

be custom built for each

mission.

Standardized payload accommodations were not

provided by the STS; therefore, all electrical,

mechanical and fluids accommodations are custom

designed for every mission. Also standardized

mission planning for payload mass, orbit

destinations, etc. were not provided; therefore, each

mission is planned as a custom mission.

There were no structured engineering management

processes put in place to provide constraints or to

limit these functional requirements for each flight.

There was no automated functional verification

capability (IVHM) provided to reduce the labor

intensiveness of the task.





realistic and based on the functional requirements

along with assurance the objectives would be met.


21

9/28/2010

Attribute

I D.




Current Systems


Critical Shortfalls


R 3.6

R 3.9

A 1.5



mission

Total number of manual



critical flight functions,

e.g., CG, fluid residuals

content & purity,

functionality of primary

and backup system

hardware

Vehicle, payload, and

ground systems

integration functions


all LCC requirements by


assumptions.

Con’t


402 functional interfaces of

which the Propulsion

discipline alone has 236 of

which the SSME has




system for the single


station (ICD-2-1A002) –

for the vehicle to ground

design and operations

activities. Note: The orbiter

element has ten (10) more

facility station ICD’s at the

launch site.

The above is an example of

all major flight element

interface support

requirements as the SRB’s

have 16 safety driven

functional requirements

and 28 safety driven

limited access control

requirements.


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R 3.1

&3.8

12. Mission Planning Cycle

Was considered

within the 40

flights/yr. with 4

vehicle fleet and

the 24 hr. notice

to launch

requirement.

400 day typical cycle Standardized payload accommodations were not


mechanical and fluids accommodations are custom

designed for every mission. Also standardized

mission planning for payload mass, orbit

destinations, etc. were not provided; therefore, each

mission is planned as a custom mission.


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4.2 Future STS Needs

The pursued US Exploration Program must be "sustainable" (i.e., it must be within

budget and within yearly budget caps both during procurement and throughout its long

operating life). For this to be achievable, operability must be designed into the

architectures and elements from the very beginning. Indeed, NASA is attempting to

implement this as shown by NASA NPR: 7120.5C (This document is a “Must Read”)

dated February, 2005. NASA Program and Project Management Processes and

Requirements, Paragraph 6.2.3 Systems Engineering Requirements: The Project

Manager and project team shall:

(a.) With the Program Manager, customers, and stakeholders, define a validated

set of Level 1 requirements and success criteria for the project in Phase A.

(b.) Develop operations scenarios and concepts, mission profiles, and mission

operational modes for the purpose of fostering a better understanding of operational

requirements, including LCC drivers for logistics and maintenance.

To further this effort the Space Propulsion Synergy Team has developed, over a number

of years and a number of separate tasks, a series of Technical Performance Metrics

(TPMs) or control needs that would help assure a sustainable operational space

transportation system architecture. The following section summarizes these TPMs.

How to Improve the Control of Life Cycle Cost (LCC)

The following are a recommended listing of ―Design for Operability‖ requirements TPMs

control needs. The purpose of these ―requirements‖ is to guide and control the

development of the overall and element architectural concepts and the designs of vehicle

components, subsystems and systems in order to minimize and control LCC by focusing

on operations and maintenance costs drivers. These needs are a response to this shortfalls

analysis performed on the Shuttle program and reflect the major lessons learned.

It is suggested that a listing of those focus-area measurable criteria that require an

Engineering Management structured process be established within the requirements

documentation are as follows:

1. Total number of separate identified vehicle propulsion systems (lack of discipline

functional integration). This also applies to number of separate stages: Metric Value:

TBD

Shuttle Reference Value: Many systems in MPS, OMS, RCS, TVC, Thermal

Management Systems and Life Support Systems

2. Total number of flight tanks in the architecture: Metric Value: TBD

Shuttle Reference Value: In excess of 95

3. Number of safety driven functional requirements to maintain safe control of systems

during flight and ground operations: Metric Value: TBD


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9/28/2010

Shuttle Reference Value: In excess of 70

4. Number of maintenance actions unplanned between missions: Metric Value: TBD

Shuttle Reference Value: ~ 800

5. Number of maintenance actions planned between missions: Metric: Value TBD

Shuttle Reference Value: ~ 2200

6. Total number of traditional ground interface functions required: Metric Value: TBD

Shuttle Reference Value: Hundreds

7. Percent (%) of all systems automated: Metric Value: TBD

Shuttle Reference Value: (Inspections and checkout mostly manual)

8. Number of different fluids required: Metric Value: TBD

Shuttle Reference Value: 24 every flight

9. Total number of vehicle element to element support systems (Major element

interfaces such as Orbiter to SSME or ET): Metric Value: TBD

Shuttle Reference Value: Example is the SSME with 26 support systems from the

Orbiter (target value should be 12 or less)

10. Number of flight vehicle servicing interfaces: Metric Value: TBD

Shuttle Reference Value: ~102

11. Number of confined/closed compartments: Metric Value: TBD

Shuttle Reference Value: 13 or more

12. Number of commodities used that require medical support operations and routine

training: Metric Value: 0 Toxics & TBD Special Training

Shuttle Reference Value: 3 major & 3 minor toxic fluids

13. Number of safety driven limited access control operations: Metric Value: TBD

Shuttle Reference Value: In excess of 266 functions

14. Number of safing operations at landing: Metric Value: TBD

Shuttle Reference Value: TBD

15. Number of mechanical element mating operations (element to element & element to

ground): Metric Value: TBD

Shuttle Reference Value: Example: 24 component mating between the one

SSME and the Orbiter (A total of 72 total SSME mechanical connections to the

Orbiter) Target for a single engine to stage should be more like 9 to 11.


25

9/28/2010

16. Number of separate electrical supply interfaces: Metric Value: TBD

Shuttle Reference Value: Example: 12 electrical components matings needed for

each SSME to the Orbiter (A total of 36 total SSME electrical connections to the

Orbiter) Target for a single engine to stage should be 4 or less.

17. Number of intrusive data gathering devices: Metric Value: TBD

Shuttle Reference Value: Example : 45 intrusive sensors on each SSME

18. Number of Criticality – 1 (Crit-1) system and failure analysis modes: Metric Value:

TBD

Shuttle Reference Value: Example is that there are 550 Crit 1 & 1R failure

modes on each SSME

Reference cases 1 and 2 of data for the Shuttle in enclosed as Appendices V and VI.

4.3 Fluid & Propulsion System Technical Generic Needs for Future STS (flight &

ground systems)

The Fluid & Propulsion System Technologies Focused Needs for Future Flight

and Ground Space Transportation Systems are determined by the desired

characteristics focus as follows:

1. Non-intrusive instrumentation

2. Process instrumentation

3. Zero emissions components

4. Elimination of the need for dynamic seals

5. Integrated functions to minimize total number of systems

6. Select passive solutions vs. dynamic components and systems

7. Reduction of total parts count by orders-of-magnitude

8. Increased design life by orders-of-magnitude

9. Increase the Reliability of components and systems by orders-of-

magnitude

10. Balance the Maintainability, Safety, and Reliability requirements to

produce control of recurring cost by design

11. Produce solutions with orders-of-magnitude less waste

12. Provide smart hardware to provide for health management and control

13. Fully automated systems that result in order-of-magnitude net reductions

of personnel requirements


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27

9/28/2010

4.4 Balancing Safety, Reliability, and Maintainability Requirements

A process and methodology for developing and balancing the quantitative SR&M requirements

have been developed at NASA KSC by Tim Adams and Russel Rhodes for the SPST.

The process for developing and balancing quantitative requirements for safety (S), reliability (R),

and maintainability (M) is shown in Figure A. This process derives and integrates Level I

requirements and the controls needed to obtain Program key objectives for Safety and Recurring

Cost.

The process being quantitative, uses common and standard mathematical models. Even though

the process is shown as being worked from the top down, this process can be worked from the

bottom up. Figure B provides two illustrations using this process.

This process uses three math models. Starting at the top, the math models are the Binomial

Distribution (greater than or equal to case), reliability for a series system, and the Poisson

Distribution (less than or equal to case). The Binomial Distribution is equivalent to the

commonly known Exponential Distribution or ―constant failure rate‖ distribution. Either model

can be used; the Binomial Distribution was selected for modeling flexibility since it conveniently

addresses both the zero fail and failure cases. The failure case is typically used for non-human

occupied spacecraft as with missiles.

As the first step of the process, the Systems Engineering Designer begins with three inputs,

namely, the desired number of missions the program is planning (n), the minimum number of

successful missions for duration of the program (x), and assurance (A) of obtaining x or more

successes out of the n missions. In risk terms, 1 - A is the probability or risk likelihood of not

obtaining x or more successes out of n number of attempts or not obtaining the desired level of

safety and reliability. When these three mentioned inputs are used in the Binomial Distribution,

the minimum mission reliability (Ps) is calculated. At this point of the process, the Level I

Safety requirement has been established.

The second step uses the minimum mission reliability (Ps) and an estimate of the number of

serial LRU elements (e) as inputs into the formula for reliability of a series system to calculate

minimum element reliability (Psi). Maximum element failure rate (Pfi) is equal to 1 - Psi.

Without considering the maintainability burden that has a very large influence on recurring cost,

the process at this point has established the Safety and Reliability requirements for the program.

The last step addresses the maintainability parameter, the parameter that provides a control for

recurring costs due to maintenance and repair. Similar to program reliability (A), program

maintainability (M) is a probability. The probability M is determined by the Poisson

Distribution and uses the following inputs: the number of missions (n), the number of elements

(N, where e N), the LRU failure rate (Pfi or λ, where λ Pfi), and the maximum number of

LRU repairs (r). Technically, M is the probability of no more than r number of repairs

occurring at a particular mission using e number of LRU’s with an average failure rate of Pfi or

λ. To achieve the desired results in both M and the desired A, adjustments in e, Pfi, N, and λ

must be made. These values become the enabling requirements to balance and achieve the

desired key objectives of the program.


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Number Of Number Of Probability ( A ) Of x Or More

Missions ( n ) Successful Missions ( x ) Successes Out Of n Missions

Note: 1 - A = Program Risk

Minimum Mission

Reliability ( Ps )

Number Of Serial System Elements ( e )

Minimum Reliability

For Any Element ( Psi )

Maximum Failure Rate

For Any Element ( Pfi ) Note: 1 – Psi = Pfi

Maximum Number Of

System Elements Expecting

Repair ( r ) Per Mission

Probability ( M ) Of No More Than

r Number Of Repairs At Mission n

Figure A

Via Math Model

Via Math Model

Via Math Model

M

Maintainability

R

Reliability

S

Safety

Repair Rate ( λ ),

use Pfi when e = N

Repairable

Elements ( N ),

use e when e = N


29

9/28/2010

S & R Table: Maximum Failure Rate For Each Serial System (Pfi)

Number of Serial Systems (e)

10+1

10+2

10+3

10+4

10+5

Ass

ura

nce

(A

)*, th

e

pro

bab

ilit

y o

f 10

0

succ

esse

s ou

t of

100

att

emp

ts

.90 1.0535 x 10-4

1.0536 x 10-5

1.0536 x 10-6

1.0536 x 10-7

1.0536 x 10-8

.95 5.1292 x 10-5

5.1293 x 10-6

5.1293 x 10-7

5.1293 x 10-8

5.1293 x 10-9

.99 1.0050 x 10-5

1.0050 x 10-6

1.0050 x 10-7

1.0050 x 10-8

1.0050 x 10-9

.999 1.0005 x 10-6

1.0005 x 10-7

1.0005 x 10-8

1.0005 x 10-9

1.0005 x 10-10

Note: * Assurance (A) is a composite of safety (S) and reliability (R).

M Table: The Probability (M) Of No More Than 1 Element Repair Per Mission

Number Of Subsystem Elements** (N) At The Repair/Maintenance Level

10+2

10+3

10+4

10+5

10+6

Ma

xim

um

Rep

air

Ra

te F

or E

ach

Ele

men

t (λ

)

10-3

0.9953

0.7356

0.0005

0

0

10-4

0.99995

0.9953

0.7356

0.0005

0

10-5

0.9999995

0.99995

0.9953

0.7356

0.0005

10-6

0.999999995

0.9999995

0.99995

0.9953

0.7356

Note: ** When necessary, count legs in a redundant system as subsystem elements.

Illustration 1 (e = N case):

If A = 0.99 for 100 successes out of 100 attempts and e = 100, then Pfi 1 x 10-6

will satisfy

the Assurance Requirement. Since N = e = 100 and Pfi = λ = 1 x 10-6

, then the probability of

having more than one repair per mission is remote (1 – M = 1 – .999999995 = 5 x 10-9

). Thus,

the Maintainability Requirement at virtually any level will be satisfied.

Illustration 2 (e < N case):

If A = 0.99 and e = 100, then Pfi 1 x 10-6

will satisfy the Assurance Requirement. Assume

each of the 100 serial systems contain an average of 1,000 sub-elements, then N = 1,000 x e =

1,000 x 100 = 100,000. Also, assuming each sub-element repair rate is λ = 1 x 10-5

and with N

= 100,000, then the probability of having no more than one repair per mission is 0.7356 or

about 74% -- other words, a 26% chance of 2 or more (up to N) repairs. Thus, the

Maintainability Requirement at a selected 90% level will not be satisfied.

Figure B


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4.5 A Generic Functional Systems Breakdown Structure (SBS) for Space

Transportation Architectures

A Functional SBS is a method that will provide a successful framework for defining and

specifying the requirements and can also be used for determining the general support

infrastructure needs. It also can serve as a guide for insuring LCC assessments have full

accountability of all functions required.

A generic functional SBS provides a universal hierarchy of required space transportation

operational functions, which include ground and space operations as well as infrastructure. The

matrix provides a structured, indentured breakdown of Systems’ Functional System

Requirements for the use in design definition and accountability for all functions; i.e., a giant

check list to be sure that no functions are omitted especially in the early architectural design

phases.

The Functional SBS furnishes inputs for analysis of any concept and provides a systematic

source for determining and documenting the requirements and the ―Life Cycle Costs‖ necessary

to achieve the Program/Project goals and objectives. When used correctly, the Functional SBS

furnishes a framework for defining requirements, which will prevent over or under specifying

these requirements.

This Functional SBS provides inputs for analysis of concepts and provides a source for

determining and documenting requirements necessary to achieve full accountability of Top Level

Goals. This Functional SBS will also serve as a guide to assure that the required skills are

available to support the program’s needs.

5.0 Summary

The shuttle shortfalls assessment by the national SPST provides insight into the major areas that

needs improvement as well as to the kind of operational criteria that needs to be addressed. This

assessment along with other supporting analysis provides a high potential for LCC cost reduction

and control by developing and implementing a set of proposed operability design requirements,

e.g., technical performance metrics (TPMs). This shuttle shortfalls analysis provides the insight

that a structured engineering management process would require to budget and control the TPMs

throughout the entire concept to DDT&E completion phases of any future program for LCC

controls needed to attain a sustainable NASA exploration program.

The objective of this report is to assure that the planning and implementation of the

transportation systems required by the Space Exploration Program takes maximum advantage of

the ―lessons learned‖ from the major space programs of the past decades. The focus of this report

is on what has been learned about the assessment and improving control of Life Cycle Costs

(LCC) from major space programs. The major ―lesson learned‖ from these studies is that much

improved, innovative processes must be developed and rigorously applied to effectively control

life cycle cost.


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9/28/2010

The only shuttle program major objective that was controlled with the use of a structured

Engineering Management process was performance closure by managing all flight systems

weight. Objectives were set for Life Cycle Cost (LCC) for the Shuttle, but no Engineering

Management processes were exercised to provide control (only the DDT&E cost was tracked).

These LCC objectives are of the same importance as placing a mass in orbit and must all be

managed by the same level of discipline. The NASA must do better to achieve the Presidential

requirement of conducting/achieving a sustainable space exploration program. The major portion

of the space transportation LCC is the recurring or operational phase cost.

This space shuttle shortfalls assessment study and its supporting analysis provide a major source

of documented knowledge of the shortfalls that developed between initial

requirements/objectives and the actual results achieved during the Shuttle Program. The results

of this study are included in this report. A major ―lesson learned‖ from these activities is the

importance of first clearly defining, flowing down, and controlling the ―systems requirements‖

and maintaining control throughout the R&D Program. The SPST has emphasized the need to

clearly define the ―requirements‖ up front: that is the ―what’s‖ required of the desired space

transportation system. To sum up this lesson learned, we must change the way we do business to

avoid ―doing what we always do and achieving what we always got‖. Therefore, we must change

our Engineering Management processes to include a structured process to control those major

operational functions that are major cost influences to provide the LCC controls required for a

sustainable Space Exploration Program.

Recently the SPST developed a new approach for formulating ―requirements‖ that will provide

full accountability of all functions required to perform the planned space missions. The approach

as described in this report developed a top-level functional systems breakdown structure,

(Functional SBS) with modular sub sets, that may be utilized as a basis for defining the desired

―functional requirements‖ in any space system. This process is intended to serve as a guide in

development of the work breakdown structure (WBS), provide visibility of those technologies

that need to be developed to cover a required function, and help identify the personnel skills

required to develop and operate the space transportation system for this very large and

challenging National effort. This Functional SBS covers all transportation elements on earth, the

moon and mars including any orbiting operational space nodes if deemed necessary.

Another study performed by the SPST was a ―bottom-up‖ analysis as to why past programs were

not achieving the desired functional criteria: ―What has impeded or prevented the application of

good systems engineering and management’s successful implementation of the

approaches/processes addressed in this report?‖ It was found that there are several reasons for the

impediments: lack of overall integration (stove-piping or optimizing at the single function level),

inappropriate starting technology level, the lack of sufficient Engineering Management

processes, and that many of the systems engineering requirements (needs), were ―boring‖ not

stimulating (not sexy). This indicates that major improvements in discipline must be rigorously

imposed on the system engineering and design processes by the program managers.

The desired thrust resulting from this effort is for the NASA to respond to these insights gained

in the analysis/studies referred to in this report and focus on developing the needed engineering


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management processes that will be required for NASA to achieve a sustainable space exploration

program by controlling the space transportation system’s LCC.

6.0 Conclusions

Based on the study and analysis of several space programs including the Space Shuttle by

the SPST, it is clear that past and current efforts to control life cycle costs have been

inadequate and ineffective.

The ―lesson learned‖ from these studies is that much improved, innovative processes

must be developed and rigorously applied to adequately control life cycle costs. These

improved/innovative process need to be enforced by the Program Managers throughout

the design development, production and operation of the space systems that will be

required for the Space Exploration Initiative missions. Additionally, the Safety,

Reliability, and Maintainability requirements must be balanced to provide the necessary

controls on the maintenance burden and frequency of the depot cycle to gain control of

process flow time and recurring cost.

It is believed that improved life cycle cost control processes developed by the SPST will

provide the necessary cost controls when properly applied in the future to advanced space

transportation systems.

7.0 Recommendations

It is clear that past and current efforts to control life cycle costs have been inadequate and

ineffective; therefore, the SPST recommends the NASA consider adopting the proven methods

of controlling weight and performance and applying them to controlling cost. It is also

recommended that the NASA consider using the SPST technique for balancing Safety,

Reliability, and Maintainability requirements to provide controls on recurring maintenance

burden to provide operational effectiveness and LCC.

The SPST recommends that the Program Analysis and Evaluation Office endorse these

SPST recommendations and implement requirements for all new NASA programs to

require these LCC and operational controls. Further the SPST recommends that these

new approaches be implemented immediately within the current planning of the Space

Exploration Program Missions.

We emphasize these recommendations, because, the Space Exploration Program must not

only be ―affordable‖ but ―sustainable‖. This requires close control of life cycle costs

within established budgets.

The final recommendation is for the SPST Functional Requirements Sub-team to develop

an approach to providing LCC controls on major operational cost drivers and provide it to

the NASA for their implementation consideration in the Space Exploration Program.


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Appendix I Space Shuttle Level I and II Program Requirements Documentation excerpts U.S. Gov't

National Aeronautics and Space Administration

SPACE SHUTTLE

PROGRAM REQUIREMENTS DOCUMENT

Revision NO. 8

LEVEL I

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION OFFICE OF SPACE FLIGHT WASHINGTON, D.C. 20546

June 30, 1977


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1.0 INTRODUCTION. 1.1 PURPOSE AND SCOPE. The purpose of this document is to establish the Level

I program requirements for the Space Shuttle Program. These are requirements established by the Director of the Space Shuttle Program as necessary to achieve the objective of the Space Shuttle Program, namely to: (a) reduce substantially the cost of space operations, and (b) provide a capability designed to support a wide range of scientific, defense, and commercial uses. All Space Shuttle Program planning and direction of NASA Centers should be in accord with the requirements stated herein unless specific exception is approved in writing as an addendum to those Space Shuttle requirements by the Director of the Space Shuttle Program. 1.2 CHANGES. This document will be controlled in accordance with approved Space Shuttle Program Directive No. 1. 1.3 RELATED DOCUMENTS. This document is in accord with the approved program approval document and program plan. Further detail pertaining to technical and operational requirements and to payload accommodations can be found in Level II documentation. 2.0 SPACE SHUTTLE SYSTEM REQUIREMENTS. 2.1 DESCRIPTION. The Space Shuttle System flight hardware shall consist of a reusable orbiter Vehicle including installed main engines, an expendable External Tank and reusable Solid Rocket Boosters which burn in parallel with the main engines. The Orbiter Vehicle shall be capable of crossrange maneuvering during entry, aerodynamic flight and horizontal landing. 2.2 OPERATING LIFE. As a design objective, the Orbiter Vehicle should be capable of use for a minimum of 10 years, and capable of low cost refurbishment and maintenance for as many as 500 reuses. 2.3 PAYLOAD BAY GEOMETRY. The payload bay shall be sized to have a clear volume of 15 ft. (4.5 meters) diameter by 60 ft. (18.2 meters) length. Payloads including their thermal and dynamic deflections shall be contained in an envelope equal to or less than 15 ft. (4.5 meters) in diameter and 60 ft. (18.2 meters) length. Payload attachment fittings and umbilicals shall extend beyond this envelope in order to mate with standard orbiter fittings which are outside the payload envelope. A standard deployment mechanism and tie points shall be chargeable to the Orbiter Vehicle and shall not occupy the clear volume when stowed. Clearance for deployment and Orbiter deflections shall be provided by the Orbiter Vehicle. Available payload volume is reduced when the orbiter Maneuvering System (OMS) incremental Delta V tankage or the docking module is carried.

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,U.S. Gov't

2.4 PAYLOAD MASS ACCOMMODATION. The Space Shuttle System shall be capable of operating within the up payload range from zero to 65,000 lbs. (29,483 kg) for nominal launches and abort modes. Nominal down payloads shall be limited to 32,000 lbs. (14,515 kg). The Orbiter Vehicle payload C.G. limits for longitudinal, vertical and lateral axes are shown in Figures 2-1, 2-2 and 2-3. 2.5 CREW/PASSENGER ACCOMMODATIONS. The cabin shall be designed to accommo0ate a total crew of seven, three crewmen to operate the orbiter and up to four payload specialists. The orbiter shall be provisioned for support of these personnel for 28 man days and up to 42 man days with no system change. All crew systems (such as seats and intercoms) for crew size greater than four and all consumables for duration greater than 28 man days shall be provided in kit form and shall be charged to payload. The design shall not preclude installation of crew support equipment for a total of 10 crew members as would be required to implement an Orbiter-to-Orbiter rescue. 2.6 CABIN ATMOSPHERE. The Orbiter crew and passenger environment shall be a shirt-sleeve, nominal 14.7 psi (760 mm Hg), two gas atmosphere (Nitrogen-Oxygen) to simulate sea level composition. 2.7 EXTRA VEHICULAR (EVA) PROVISIONS. The orbiter shall provide an airlock, crew provisions and support hardware for crew access to and from the unpressurized payload bay and pressurized modules, for orbiter and payload EVA operations, and for space rescue. The airlock will be capable of being mounted either inside or outside the cabin on the forward bulkhead of the payload bay and will be capable of being used in conjunction with a spacelab tunnel adapter to provide continuous spacelab-to-cabin access during EVA. To support rescue, all Shuttle Flights will carry EVA provisions for two trained crewmen and personnel rescue systems for all other crew members. # 2.8 REDUNDANCY. The redundancy requirements for all flight vehicle subsystems (except primary structure, thermal protection system, and pressure vessels) shall be established on an individual subsystems basis, but shall not be less than fail-safe. "Fail-safe,, is defined as the ability to sustain a failure and retain the capability to successfully terminate the mission. Redundant systems shall be designed so that their operational status can be verified during ground turnaround and to the maximum extent possible while in flight. 2.9 SPACE SHUTTLE MAIN ENGINES (SSME). The Space Shuttle Main Engines will meet the requirements specified in the approved

*Information changed by Revision 8 CHANGE NO. 5

8

# For record of Deviation/Waivers granted against this requirement, refer to JSC 07700, Volume X.

Comment [leh1]:


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Space Shuttle Orbiter Vehicle/Main Engine ICD. Three engines will be used in the orbital flight configuration. * 2.10 REACTION CONTROL SUBSYSTEM (RCS). An Orbiter RCS shall provide three-axis angular control, including vernier angular control, and three-axis translation capability. * 2.11 ORBITAL MANEUVER SUBSYSTEM (OMS). The OMS shall provide the propulsive thrust to perform final insertion into orbit, circularization, orbit transfer, rendezvous and deorbit. The OMS tankaqe shall be sized for a Delta V capability of 1,000 ft/sec (305 m/sec) with a 65,000 lb payload. This Delta V capability includes the final orbit injection Delta V from external tank separation to 50 x 100 n. mi. (93 x 185 km) insertion orbit. Provisions shall be made to allow additional tankage to be incorporated in three Delta V increments of 500 ft/sec (152 m/sec) each for an overall total Delta V capability of 2,500 ft/sec (762 m/sec). The additional tankage and propellants will be located id the payload bay and the weights and volumes thereof charged to payload. 2.12 AIRBREATHING ENGINE SUBSYSTEM (ABES). (Deleted). 2.13 SOLID ROCKET BOOSTERS (SRB's). The SRBs will meet the requirements specified in the approved Space Shuttle Orbiter Vehicle/ET/SRB Interface Control Document. The two SRBs will operate in parallel with the main engines to provide impulse to the Orbiter Vehicle from lift-off to staging. the SRBs shall be designed for water recovery, refurbishment and subsequent reuse. As a design objective, the SRB case should be capable of 20 uses. * 2.14 EXTERNAL TANK (ET). The expendable External Tank will carry hydrogen and oxygen propellant for the main engines. The ET will conform to the requirements of the approved Space Shuttle Orbiter Vehicle/ET Interface Control Document. 2.15 RADIATION AND AVIONICS. (Deleted). 2.16 COMMUNICATIONS SUBSYSTEM. The Orbiter shall be capable of direct voice command, telemetry and video communication with the ground. The Orbiter shall be capable of communication by relay through a communication satellite system. Provisions shall be made to accommodate equipment for secure voice and data communication. 2.17 LANDING SYSTEM. The Orbiter Vehicle shall have an automatic landing system. 2.18 SAFETY, RELIABILITY, MAINTAINABILITY AND QUALITY. The provisions of NHB 5300.4 (lD-1), August 1974. "Safety, Reliability, Maintainability and Quality Provisions for the Space Shuttle Program" will apply for the Space Shuttle Program.

*Information changed by Revision 8 CHANGE 2


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9 U.S. Gov't 2.19 INTERNATIONAL DOCKING SYSTEM. The Orbiter Vehicle shall meet the international requirements negotiated for compatible rendezvous, docking and crew transfer systems. The docking module will be provided as an optional kit in the payload bay and will be chargeable to payload. 2.20 ELECTROMAGNETIC COMPATIBILITY. The Space Shuttle system, shall be designed and tested in accordance with JSC Specification SL-E-0001. Subsystems and/or individual equipment shall be designed and tested in accordance with JSC specification SL-E0002. *2.21 RANGE SAFETY FLIGHT TERMINATION SYSTEM. The Shuttle vehicle shall have a range safety flight termination system for Orbital Flight Test Operations involving an Orbiter equipped with ejection seats. *2.22 FACILITIES AND SUPPORT EQUIPMENT COMMONALITY. A major goal of the Space Shuttle Program shall be to minimize the national investment in launch facilities, GSE, and other support equipment (including the launch processing system and associated software) through maximization of the commonality of requirements, design and procurement of these items between KSC and VAFB. The specification and design of operational facilities, support equipment and procedures at KSC shall include maximum consideration of the requirements and design constraints inherent in operations at VAFB. VAFB design shall make maximum practical use of the operating procedures and ground and other support equipment developed for KSC. 3.0 OPERATIONAL REQUIREMENTS. 3.1 GENERAL. The Space Shuttle System shall be designed to accomplish a wide variety of missions. The Shuttle System weight carrying capability into orbit shall be based on the performance required to execute mission 3A. The equivalent maximum performance is shown in Figures 3-1 and 3-2 for the range of inclinations and altitudes indicated. The payload capability curves assume a simple deployment mission with no rendezvous, 22 fps (6.9 m/sec) OMS Reserves, 4,500 lbs. (2,041 kg) of RCS propellant, and direct deorbit. (Reentry performance restrictions are addressed in Par 2.4. Detailed Shuttle System performance questions should be addressed to the JSC Shuttle Program office). Space Shuttle missions will involve direct delivery of payloads to specified low Earth orbits; placement of payloads and transfer stages in parking orbits for subsequent transfer to other orbits; rendezvous and station keeping with detached payloads for on-orbit checkout; return of payloads to Earth from a specified orbit; and provisions for routine and special support to space activities, such as sortie missions, rescue, repair, maintenance, servicing, assembly, disassembly and docking

*Information changed by Revision 8

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3.2 REFERENCE MISSIONS. 3.2.1 Design Reference Missions. These missions shall be used in conjunction with the other requirements specified herein to size the Space Shuttle System. For performance comparison)s, Mission 1 will be launched from Kennedy space center (KSC) into a 50 by 100 n. mi. (93xl65 km) insertion orbit, and Mission 3 will be launched into the same insertion orbit from the Vandenberg AFB.

a. Mission 1. Mission I is a payload delivery mission to a 150 n. mi. (278 km) circular orbit. The mission will be launched due east and requires a payload capability of 65,000 lbs. (29,483 kg). The Boost phase shall provide insertion into an orbit with a minimum apogee of 100 n. mi. (185 km) , as measured above the Earth's mean equatorial radius. The purpose of this mission is assumed to be placement of 65,OOC lb. (29,483 kg) satellite and/or retrieval of a 32,000 lb. (14,515 kg) satellite. The Orbiter Vehicle orbit translational Delta V requirements in excess of a 50 by 100 n. mi. (93 x l85 km) reference orbit are 650 ft/sec (198 m/sec) from the Orbital Maneuver Subsystem (OMS) and 100 ft/sec (30 m/sec) from the RCS.

b. Mission 2. (Deleted) .

c. Mission 3. Mission 3 shall consist of two missions, one for payload delivery

and one for payload retrieval. This is a 3-day, 2-man mission.

d. Mission 3(A). This mission is a payload delivery mission to an orbit of 104 degrees inclination and return to the launch site. The boost phase shall provide insertion into an orbit with a minimum apogee of 100 n. mi. (185 km) as measured above the Earth’s equatorial radius. The Orbiter Vehicle on-orbit translation Delta V requirements in excess of a 50 by 100 n. mi. (93 X 185 km) reference orbit are 250 ft/sec (76 a/sec) from the orbital Maneuver Subsystem (OMS) and 100 ft/sec (30 m/sec) from the RCS. The ascent payload requirement is 32,000 lbs. (14,515 kg). For mission performance and consumables analysis, a return payload of 2,500 lbs. (1134 kg) will be assumed (the 2500 lbs. (1134 kg) is included in the 32,000 lbs. (14,515 kg)

ascent payload weight).

e. Mission 3(B). This mission is a payload return mission from a 100 n. mi. (185 km) circular orbit. it 104 degrees inclination and return to the launch site. The

CHANGE NO. 1 11


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return payload weight is 25,000 lbs. (11,340 kg). For mission performance

and consumables analysis, an ascent payload of 2,500 lbs. (1134 kg) will be assumed (the 2,500 lbs. (1134 kg) is included in the 25,000 lbs. (11,340 kg) return payload weight). The Orbiter Vehicle on-orbit translation Delta V requirement in excess of a 100 n. mi. (185 km) circular orbit is 425 ft/sec (130 a/sec) from the OMS. The translational Delta V requirement from the RCS is 190 ft/sec (58 m/sec) .

3.2.2 Performance Reference Missions. These missions shall be used in conjunction with the other requirements specified herein to assess the performance capabilities of the Space Shuttle System, as sized by the design reference missions, to assure that the mission requirements will be met.

a. Mission 4. This mission is a payload delivery an(I retrieval mission launched from the Vandenberg AFB Launch Site to a final inclination of 96 degrees in a 150 n. mi. (277.8km) circular orbit as measured above the Earth's equatorial radius. The ascent payload requirement is 32,000 lbs. (14,525 kg). The return payload requirement is 25,000 lbs. (11,340 kg). The Orbiter vehicle on-orbit translational Delta V requirement, including post MECO insertion burn and deorbit, is a total of 1,050 ft/sec (321 m/sec). The onboard RCS propellant tanks will be fully loaded at launch.

3.3 LAUNCH AZIMUTH. The Space Shuttle System shall have a variable azimuth launch capability to satisfy the acceptable launch-to-insertion azimuths from both the KSC and Vandenberg AFB launch sites. 3.4 CROSSRANGE. The Orbiter Vehicle shall have the aerodynamic crossrange capability to return to the launch site at the end of one revolution for all inclinations within the Space Shuttle System capability. Crossrange is to be achieved during entry, which is defined as beginning at 400,000 ft (122 km) altitude and ending at 50,000 ft. (15 km) altitude. 3.5 RETURN PAYLOAD. The Orbiter Vehicle shall have the capability to land the design return payload of 32,000 lbs. (14,515 kg) with nominal wind and load factors and up to 65,000 lbs. (29,483 kg) return payloads under increased landing condition constraints. 3.6 LOAD FACTORS. The Space Shuttle System launch trajectory resultant load factors shall not exceed 3 G’s for the Orbiter Vehicle. These limits do not apply to abort modes. The product of G forces and time shall not be detrimental to the crew/passengers. 3.7 TURNAROUND. When operational the Space Shuttle System flight hardware turnaround from landing return to the launch facility to launch readiness shall not exceed 160 working hours covering a span of 14 calendar days for any class mission.

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3.8 LAUNCH FROM STANDBY. The Space Shuttle System design shall provide the capability to be launched from a standby status within 2 hours, and hold in a standby status for 24 hours. Standby status is defined as ready for launch except main propellant fill, crew ingress and final systems verification. Waiver: 1) For KSC launches, the time from standby to launch shall be a 4-hour capability. 2) For VAFB launches, the time from standby to launch shall be four (4)

hours. This may be reduced to three (3) hours should parallel crew ingress during MPS loading be permitted and to two (2) hours by further automating the cabin pressure integrity check.

3.9 RESCUE. To fulfill the space rescue role, the Space Shuttle System shall have the capability to launch within 24 hours after the vehicle is mated and ready. for transfer to the pad. if the spacecraft requiring aid has a docking system on that mission, the primary rescue mode will te ry docking, with crew transfer through a pressurized tunnel. Otherwise, emergency rescue will be with pressure suits and personal rescue systems outside the spacecraft. The Orbiter Vehicle shall be capable of supporting the survival of a 4 man crew for 96 hours after an on-orbit contingency. Support for additional personnel shall be provided by the Orbiter and charged to the payload per paragraph 2.5. Waiver: For KSC launches, the time requirement from Notification to launch shall be 26.5 hours. 3.10 ABORT. The Space Shuttle System shall provide a safe mission termination capability through all mission phases. The performance capability to meet this requirement is defined as follows:

a. Crew and Passenger Ingress Through Launch Commit Phase. The emergency egress shall provide for crew add passenger evacuation to a safe area in a maximum time of two minutes (from crew ingress to swing arm retract).

b. Launch Commit Through Return to Site Capability Phase. The Shuttle

System shall have a performance capability of intact (crew, payload and vehicle) abort and return to the launch site. The system design shall include adequate provisions for external tank separation and disposal.

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c. Return-to-Site Through Orbit Insertion Phase. The orbiter shall have the

capability (with one main engine out) to abort once around and return to the primary landing site from the point in the flight trajectory where a direct return-to-launch-site capability ends.

d. Orbital and Reentry Phase. The abort mode after orbit insertion shall be

early mission termination and return to a suitable landing site. 3.11 LOITER TIME. (Deleted). 3.12 ORBITER TRANSPORT. The Orbital Vehicle shall be capable of being transported by carrier aircraft. 3.13 MISSION DURATION. Mission duration of 7 days shall be used to size the orbiter for self sustaining lifetime (from lift-off to landing) for a crew of four in accordance with Section 2.5. The orbiter design shall not preclude the capability to extend the orbital stay time up to a total of 30 days by adding expendables. 3.14 UNMANNED FLIGHT. (Deleted). All Space Shuttle orbital Flights shall be manned. 3.15 LAUNCH RATE. (Deleted). Refer to the currently approved Program Directive on Controlled Milestones. 3.16 OPERATIONAL DATES. (Deleted). Refer to the currently approved Program Directive on Controlled Milestones. 3.17 ORBITER VEHICLE ATTITUDE CONSTRAINTS. While the payload bay doors are open the orbiter will provide heat removal from the payload up to 21,500 BTU/hr in addition to the orbiter heat load. The addition of a radiator kit will increase the orbiter heat rejection and provide heat removal from the payload up to 29,000 BTU/hr. The radiator kit is comprised of two radiator panels added to the aft section of the Orbiter payload bay doors and the weight will be chargeable to the payload. With proper water storage and thermal conditioning before pointing the payload bay in the following typical attitudes, the minimum durations that the orbiter will maintain attitude holds are:

1 4

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Without With Payload Bay Viewing Radiator Kit Radiator Kit Deep Space **>160 Hrs **>160 Hrs (Non 3 Axis Inertia-1) Deep Space, Stellar 33 Hrs 33 Hrs (3 Axis Inertial) Direct .Earth **18 Hrs **30 Hrs

Direct Solar 12 Hrs 16 Hrs (3 Axis Inertial)

**For beta angles (orbit plane relative to the solar vector) in the range of 60 degrees to 90 degrees with worst case thermal orientation (other than 3 axis inertial holds) the orbiter shall be designed for repeated cycles of a minimum 6 hours attitude hold with no attitude constraints, followed by a maximum of 3 hours thermal conditioning. The hold times shown are cumulative times for the 6 hours hold periods.

These hold times can be extended by the selection of pre-mission variables such as vehicle orientation and orbital parameters. A maximum of 12 hours of pre-entry thermal conditioning is required. 3.18 ORBITAL POINTING. The orbiter vehicle shall be capable of pointing a vector defined in th4-? navigation base-fixed axis system to day ground or celestial object with an accuracy of ±0.5 degrees. 4.0 ORBITER/PAYLOAD INTERFACE REQUIREMENTS. 4.1 PAYLOAD DEFINITION. Payloads referred to throughout this document are construed as the collective grouping of space hardware items such as: Spacelab experiments, research equipment, satellites, support modules, adapters and fueled transfer stages or equipment, into appropriate composite flight packages. For definition of Shuttle/orbiter payload accommodations, refer to Space Shuttle System Payload Accommodations Document. In the interest of maintaining minimum interface, (clean interface philosophy), payload designs should, where possible be self-sufficient systems, capable of checkout before installation in the orbiter and adaptable to standardized interface concepts jointly developed between payloads and Space Shuttle. 4.2 CHECKOUT. Payload performance testing and payload system checkout will be required prior to installation. Payload

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checkout while on the launch pad will be minimized and physical access to the payload will be limited. On-orbit status checks of the payload will be provided via the orbiter prior to release and/or retrieval. 4.3 DATA MANAGEMENT. The Orbiter shall provide standard displays and controls for monitoring the safety status of the payload. The payload shall provide to the Orbiter, at the interface, such information concerning the status or condition of the payload as is necessary to insure safe vehicle operation. Digital discrete, and analog signals shall be conditioned by the payload and supplied to the Orbiter Vehicle. Such equipment and capability shall be chargeable to the payload. Payload unique control and display accommodation with the Orbiter cabin shall be chargeable to the payload. A minimum standard interface shall be provided to exchange data for safety and payload status checks, and vehicle and operational parameters, such as navigation, guidance and control. Additional support may be feasible during certain operational modes. 4.4 PAYLOAD COMMUNICATION. The Orbiter shall provide direct and relay telemetry command and two-way voice capability with attached payloads and with released payloads. The Orbiter shall be capable of receiving and displaying limited payload data including video information and the RF downlink shall provide for relay of those limited payload data to the ground for both attached payloads and for released payloads. 4.5 PAYLOAD SAFETY. Space Shuttle payload elements shall insure elimination or control of payload design and operational hazards to the Space Shuttle System, other payloads and personnel. Safe payload operation with a minimum dependence on the Orbiter and crew for safing actions is a Space Shuttle goal. Requirements which are to be met by payloads are defined in the NASA "Safety Policy and Requirements for Payloads Using the Space Transportation System." 4.6 CONTAMINATION. The Orbiter Vehicle shall be designed to minimize the generation, introduction and accumulation of contaminants within the cabin, payload bay, and around attached payload modules. Payload and Orbiter RCS thruster exhaust shall not impinge or be reflected on deployed payloads or into the open payload bay. The total level of contamination within the payload bay from all sources shall be controlled to minimize the effects on payloads during all phases of Shuttle operations. 4.7 POWER. The Orbiter electrical power system shall provide for payload electrical energy allowance of not less than 50 kwh in the form of DC power to payloads through the Orbiter fuel cells. For missions with greater energy requirements, kits of approximately 840 kwh each will be provided outside the 15 ft. by 60 ft. (4.5xl8.2 meters) clear payload volume and will be chargeable to payload. Power supplied by the Orbiter for payload


CHANGE NO. 2

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on orbit consumption will be limited to 7 kw average and l@? peak. (maximum duration peak power levels will be limited to 15 minutes duration at no less than 3 hour intervals.) 4.8 PAYLOAD POINTING. For payload pointing purposes, the orbiter GN&C system shall be capable of interfacing with a payload supplied and payload mounted attitude sensor. Provided that the accuracy of this sensor is equivalent to that of the Orbiter IMU, the Orbiter shall be capable of pointing a Vector defined in the sensor-fixed axis system to any ground or c4,ic-st3.cil (celestial???) object with an accuracy of +/-O.5 degrees. For payload pointing requirements beyond the capability of the Orbiter, the Orbiter GN&C system shall also be capable of interfacing with a compatible payload supplied and payload mounted stability and control system. 4.9 RENDEZVOUS AND DOCKING. The Orbiter Vehicle shall have an onboard capability to rendezvous and dock with an in-plane cooperative target or a passive stabilized orbiting element displaced up to 300 n. mi. (555 km). for Orbiter Vehicle preplanned docking missions, the docking mechanism will be installed in the payload bay. The weight of the docking mechanism and associated attachment fittings shall be chargeable to the payload. 4.10 PAYLOAD ATTACHMENT. The Orbiter shall provide standard discrete attachment points for mounting payloads. These attachment points shall be located along the payload bay, to accommodate different payload lengths and to allow for random order retrieval of multiple payloads. 4.11 PAYLOAD DEPLOYMENT AND RETRIEVAL MECHANISM. The orbiter shall provide a payload deployment and retrieval mechanism which shall be stowed outside the 60 ft. (18.2 meters) length by 15 ft. (4.5 meters) diameter payload volume. This mechanism shall deploy the payload clear of the Orbiter mold line. Release of the payload from the deployment mechanism shall leave the payload and the orbiter with only small residual rates. Spin-up capability, if required, will be accomplished by the payload. For retrieval, the Deployment and Retrieval Mechanism shall interface with payloads designed for retrieval and, after attachment of the mechanism to the payload, shall align the payload in the payload bay to accommodate secure stowage of the payload. Additionally, the Payload Deployment and Retrieval Mechanism shall be capable of supporting the payload in the deployed position under the attitude stabilization and docking loads. 4.12 PAYLOAD BAY VENTS. Provisions for venting the payload bay shall be provided by the Orbiter. This vent system shall minimize the impact of venting upon the attitude control system.


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4.13 PAYLOAD BAY ACCESS. The Orbiter and launch facility shall permit access to the payload bay for payload installation, service, and removal in the Orbiter flight preparation area and on the launch pad. Access for personnel and cargo to the payload bay shall also be available through the hatch, which interfaces the Orbiter crew compartment with the payload bay. Ground access to the payload bay with be limited to the period up to 8.5 hours before launch. 4.14 PROPULSIVE STAGES. The Orbiter design shall include provisions for fill, vent, drain and dump, of liquid propellants of propulsive stages. 4.15 ACOUSTIC ENVIRONMENT. The Orbiter Vehicle payload bay interior sound pressure level shall not exceed a maximum overall of 145 dB during liftoff sequence (T = 10 secs). and 137 dB during other mission phases. The spectral frequency distribution is shown in Table 4.1. 4.16 SHUTTLE/PAYLOAD RANDOM VIBRATION. The Space Shuttle Orbiter mid-fuselage/payload bay random vibration criteria is reflected by figure 4-4. The actual vibration input to payloads will depend on transmission characteristics of mid-fuselage; payload support structure and interactions with each payload's weight, stiffness, and c.g. Payload random vibration will result largely from direct acoustic induced vibration and will be unique for each payload configuration. 4.17 ELECTROMAGNETIC COMPATIBILITY. Payloads shall be designed and tested to be compatible with the Space Shuttle System. The Space Shuttle System will be designed and tested in accordance with JSC specifications SL-E-0001 and SL-E-0002.

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Pages 19 thru 25; Figures 2-1, 2-2, 2-3, 3-1, 3-2, Table 4-1, and Figure 4-2 NOT IN


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Appendix II

Space Shuttle Launch Information Data Base Space Shuttle Launch Information Data Base

Mis

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Land

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102

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STS 1 EAFB 23 2 Days G & C GPC timing

STS 2 EAFB 23 34 Days RCS/APU

FRCS oxidizer,APU lube oil, & minor

delay for fuel cell Ox, Mux/DeMux

STS 3 White Sands 17 1 Hr. GSE GN2 purge gas heater

STS 4 EAFB 22 6/27/1982

STS 5 EAFB 22 11/11/1982

STS 6 EAFB 22 4/4/1983 74 Days SSME SSME fuel leak (cracked valve)

STS 7 EAFB 15 6/18/1983

STS 8 EAFB 22 17 Min. Weather 17 minutes 17 Min

STS 9 EAFB 17 28 Days SRBs

Roll back & destack for nozzle

replacements

STS 41B KSC 15 2/3/1984 5 Days APU 3 APU's replaced precautionary

STS 41C EAFB 17 4/6/1984

STS 41D EAFB 17 2 Days

64 Days

GPC/MEC

GPC, GPC/SSME's, MEC for SRB

firing command, & Range 6min-50

sec aircraft

STS 41G KSC 33 10/5/1984

STS 51A KSC 15 1 Day Weather Upper wind shear 1 Day

STS 51C KSC 15 1 Day Weather Freezing cold 1 Day

STS 51D KSC 33 55 Min. Range Ship in SRB recovery zone

STS 51B EAFB 17 2Min18Sec GSE LPS failure

STS 51G EAFB 23 6/17/1985

STS 51F EAFB 23 17 Days SSME

SSME Coolant valve failure, Also

one SSME shutdown in flight @

5min.45sec. with ATO

STS 51I EAFB 23 3 Days

Weather,

GPC, &

Weather

Thunderstorms @ LC, #5GPC failed,

&Weather with Range ship in SRB

recovery area

3 Days

STS 51J EAFB 23

22 min. 30

sec. MPS LH2 prevalve issue

STS 61A EAFB 17 10/30/1985

STS 61B EAFB 22 11/26/1985

STS 61C EAFB 22 25 Days

SRB,GSE,

Weather,

GSE,

Weather

SRB HPU, LO2 sys, weather @ TAL

Sites, LO2 sys, Weather for heavy

rain

??????

STS 51L N/A 6 Days

Weather,

GSE,

Weather,

GSE

Weather @ TAL site, GSE orbiter

hatch tool stuck, Weather @ SLF

cross-wind, GSE HIM

??????

STS 26 EAFB 17 1 Hr.38 Min.

Crew sys,

Weather

Crew suit cooling fuses & upper

winds

1 Hr & 38

Min

STS 27 EAFB 17 1 Day Weather Weather visibility & winds 1 Day

STS 29 EAFB 22 1 Hr.50Min. Weather Weather Fog & upper wind

1 Hr & 50

Min

STS 30 EAFB 22 6 Days

MPS,

Weather

MPS LH2 recirc pump & leak

@ET/Orb disconnect and Weather

@ SLF/ visibility & wind

????

STS 28 EAFB 17 8/8/1989


SSME,

WeatherSSME Controller & Weather @SLF ?????

STS 33 EAFB 4 2 Days SRB Intg. Electronics assemblies

STS 32 EAFB 22 1 Day Weather Weather @ LS 1 Day ??


Crew,

Weather,

Range,

Weather

Crew sick, weather bad, Range

computer & weather again ????

STS 31 EAFB 22 14 Days APU APU replaced

STS 41 EAFB 22 10/6/1990

STS 38 KSC 33 107+ Days

MPS

LH2 Leak @ ET/Orb disconnect

required roll-back destack & received

hail damage in transfer & handling

damage


47

9/28/2010

Space Shuttle Launch Information Data Base Continued Space Shuttle Launch Information Data Base

Mis

sio

n #

Land

ing L

ocation

Orb

iter

102

launche

d o

n tim

e

Orb

iter

102 launch

dela

y tim

e

Orb

iter

102 launch

abort

s

Orb

iter

099

launche

d o

n tim

e

Orb

iter

099 launch

dela

y tim

e

Orb

iter

099 launch

abort

s

Orb

iter

103

launche

d o

n tim

e

Orb

iter

103 launch

dela

y tim

e

Orb

iter

103 launch

abort

s

Orb

iter

104

launche

d o

n tim

e

Orb

iter

104 launch

dela

y tim

e

Orb

iter

104 launch

abort

s

Orb

iter

105

launche

d o

n tim

e

Orb

iter

105 launch

dela

y tim

e

Orb

iter

105 launch

abort

s

Majo

r S

yste

m o

r F

unction c

aused

dela

y

Su

b-s

yste

m

respo

nsib

le fo

r dela

y

Laun

ch D

ela

ys

from

Wea

ther

@

Laun

ch S

ite

Laun

ch D

ela

ys

from

Wea

ther

@

Alt. Land

ing S

ite

or

Mis

sio

n C

ontr

ol


GSE,

MPS,

Payload,

MPS,

Weather

LH2 leak in GSE, LH2 leak @

ET/Orb disconnect, payload

hardware failure, LH2 leak in MPS

prevalve cover seal, then 21min. for

weather on launch day

??????

STS 37 EAFB 33 4/5/1991

STS 39 KSC 15 50 Days

Orb.Mech,

SSME

Roll-back for Orbiter ET Door hinge

replacement, SSME HPOTP

transducer & cable replacement


MPS,

GPC,

Hyd/OMS/

RCS, &

G&C

MPS transducer failure concern,

Mux/De-Mux failure that controls

OMS/RCS/Hydraulic functions, &

IMU failure


Ele Sys.,

SSME,

Cabin

Press,

Weather

@ SLF

ET/Orb Sep electronic sys. Failure,

SSME controler failure, cabin press

valve concern & weather @ SLF for

RTLS

??????

STS 48 EAFB 22 14 Min.

Gr. Com

Faulty Ground communication link

between Launch site and Mission

Control


Payload,

GSE

Failed IMU in Upper Stage, 13 Min

delay for LO2 Repl valve repair

STS 42 EAFB 22 1 Hr. Weather Weather 1 HR

STS 45 KSC 33 1 Day MPS LO2 & LH2 leakage in Orb.Aft

STS 49 EAFB 22 34 Min Weather Weather @ TAL site 34 Min

STS 50 KSC 33 5 Min. Weather Weather 5 Min

STS 46 KSC 33 7/31/1992

STS 47 KSC 33 9/12/1992

STS 52 KSC 33 1 Hr.53Min.Weather

Weather @ SLF (wind) & TAL

(visibility)

1 Hr & 53

Min

STS 53 EAFB 22 1 Hr. 25Min.Weather ET ice buildup reduction

1 Hr & 25

Min

STS 54 KSC 33 7 Min. Weather Upper Atm. Winds 7 Min

STS 56 KSC 33 2 DaysMPS Orb MPS HiPt bleed valve indication

STS 55 EAFB 22 35 DaysSSME

SSME HiPress Turbopumps obs tip-

seal retainers

STS 57 KSC 33 6/21/1993


GSE,

SRB, &

SSME

Hold down & umb. PIC failure, SRB-

HPU failure, Abort from SSME faulty

fuel sensor


Weather,

Range

Delayed 2 Hrs for bad weather then

scrubed for Range safety failure2 Hrs

STS 61 KSC 33 1 Day Weather Weather @ SLF 1 Day

STS 60 KSC 15 2/3/1994

STS 62 KSC 33 3/4/1994

STS 59 EAFB 22 4/9/1994

STS 65 KSC 33 7/8/1994

STS 64 EAFB 4 9/9/1994

STS 68 EAFB 22 43 Days SSME SSME HPOT temperature redline

STS 66 EAFB 22 11/3/1994

STS 63 KSC 15 1 Day G&C G&C IMU failure

STS 67 EAFB 22 3/2/1995

STS 71 KSC 15 3 Days Weather Weather for Launch 3 Days

STS 70 KSC 33 55 Sec. ET ET Range Safety Reciever for SRB's

STS 69 KSC 33 7 Days Ele Power Fuel Cell failure


SSME,

WeatherSSME Lox duct thickness concern ????

STS 74 KSC 33 1 Day Weather Weather @ TAL site 1 Day

STS 72 KSC 15 1/11/1996

STS 75 KSC 33 2/22/1996

STS 76 EAFB 22 3/22/1996

STS 77 KSC 33 5/19/1996

STS 78 KSC 33 6/20/1996

STS 79 KSC 15 9/16/1996

STS 80 KSC 2 Min.

MPS,

SSME LH2 leakage in aft comp.


48

9/28/2010


Mis

sio

n #

Land

ing L

ocation

Orb

iter

102

launche

d o

n tim

e

Orb

iter

102 launch

dela

y tim

e

Orb

iter

102 launch

abort

s

Orb

iter

099

launche

d o

n tim

e

Orb

iter

099 launch

dela

y tim

e

Orb

iter

099 launch

abort

s

Orb

iter

103

launche

d o

n tim

e

Orb

iter

103 launch

dela

y tim

e

Orb

iter

103 launch

abort

s

Orb

iter

104

launche

d o

n tim

e

Orb

iter

104 launch

dela

y tim

e

Orb

iter

104 launch

abort

s

Orb

iter

105

launche

d o

n tim

e

Orb

iter

105 launch

dela

y tim

e

Orb

iter

105 launch

abort

s

Majo

r S

yste

m o

r F

unction c

aused

dela

y

Su

b-s

yste

m

respo

nsib

le fo

r dela

y

Laun

ch D

ela

ys

from

Wea

ther

@

Laun

ch S

ite

Laun

ch D

ela

ys

from

Wea

ther

@

Alt. Land

ing S

ite

or

Mis

sio

n C

ontr

ol

STS 81 KSC 33 1/12/1997

STS 82 KSC 33 2/11/1997

STS 83 KSC 15 20 Min.

Cabin

Pressure

Crew hatch seal leak required

replacement

STS 84 KSC 33 5/15/1997

STS 94 KSC33 7/1/1997

STS 85 KSC 33 8/7/1997

STS 86 KSC 15 9/25/1997

STS 87 KSC 33 11/19/1997

STS 89 KSC 15 1/22/1998

STS 90 KSC 33 4/17/1998

STS 91 KSC 6/2/1998

STS 95 KSC 33 10/29/1998

STS 88 KSC 15 1 Day Orb Hydr. Hydraulic System pressure low

STS 96 KSC 15 5/27/1999

STS 93 KSC 33 7/23/1999

STS 103 KSC 33 1 Day Weather Weather 1 Day


Weather,

MEC Weather, MEC????

STS 101 KSC 15 25 Days Weather

Weather @ SLF, Weather @ SLS &

Pad, Weather @ TAL25 Days ???

STS 106 KSC 15 9/8/2000


MPS,

Weather,

GSE

MPS Pogo valve failure, Weather @

pad, FOD on GSE?????

STS 97 KSC 15 11/30/2000

STS 98 EAFB 22 2/7/2001

STS 102 KSC 15 3/8/2001

STS 100 EAFB 22 4/19/2001

STS 104 KSC 15 7/12/2001

STS 105 KSC 15 8/10/2001

STS 108 KSC 15 1 Day Weather Weather @ Launch site 1 Day

STS 109 KSC 33 1 Day Weather Weather @ Launch site too cold 1 Day

STS 110 KSC 33 4 Days GSE LH2 sys leak in vent sys


Weather,

OMS Weather , OMS GN2 sys failure???

STS 112 KSC 5 Days Weather Weather @ JSC 5 Days


Orb GO2

sys,

Weather

Orb. GO2 sys hose leak in mid-body,

Weather @TAL site???

STS 107 N/A 1/16/2003

Totals:61 flights landed @

primary landing site

17 Launch

on time

276 Days

+ 3 Hrs 20

Min.

60 Days

from 2

Aborts

6 launch

on time

85 Days +

19 Min 18

Sec

17 Days

from 1

Abort

18 launch

on time

83 Days +

7 Hrs. 2

Min. 55

Sec

121 Days

from 2

Aborts

13 launch on

time

174 Days

+ 2 Min 30

Sec

6 Days

from 1

Abort

11 launch

on time

39 Days +

41 Min

43 Days

from 1

Abort

102 total Delays = 40 Weather + 46

Veh technical + 11 GSE + 3 Range +

2 Payload delays

??/65

Missions

??/65

Missions

Grand Total

Information

61 of 113 flights

landed @ primary

landing site, or 50

required transpot

back to launch site

@ ~ $1 Million each.

* 17 of 28

or 60.7%

were on-

time

launched:

Fleet

avg=57.7%

* 6 of 10

or 60.0%

were on-

time

launched:

Fleet

avg=57.7

* 18 of 30

or 60%

were on-

time

launched:

Fleet

avg=57.7%

* 13 of 26

or 50%

were on-

time

launched:

Fleet

avg=57.7%

* 11 of 19

or 57.9%

were on-

time

launched:

Fleet

avg=57.7%

* If Shuttle launch delay is less than one day (does not result in a scrub-turnaround), it is being documented as a launch-on-time in this analysis

Therefore: Fleet Ao = OV-102 Ao + OV-099 Ao + OV-103 Ao + OV-104 Ao + OV-105 Ao / 5 Vehicles = 92.3% +

87.5% + 95.7% + 95.3% + 96.8% / 5 = 93.5% Fleet Availability

Total Shuttle fleet Workdays of 15,831 - 904 lost days from delays (caused by scrub recycle) = 14,927 divided by the

112 flights yields an average ground processing flow of 133.3 Workdays per flight. Therefore, with an average MTBF

for the fleet of 301.2

OV-099 Ao = MTBF/MTBF + MTTR

= 178.3 / 178.3 + 25.5 = 87.5%


= 377.1 / 377.1 + 17 = 95.7%


= 277.8 / 277.8 + 13.8 = 95.3%


= 309.2 / 309.2 + 10.3 = 96.8%

Of 113 mission launch attempts, 65 Missions were launched the day scheduled or 57.5% and 13 of the 48 launch

scrubs were weather related or 27%

OV-102 = 82 total

Work Days launch

delay in 8 events or

an avg. of 10.3 work-

days as its mean time

to repair (MTTR)


= 363.4 / 363.4 + 30.5 = 92.3%

OV-102 = 336 total

Work Days launch




to repair (MTTR)

OV-099 = 102 total

Work Days launch




to repair (MTTR)

OV-103 = 204 total

Work Days launch


an avg. of 17 work-


to repair (MTTR)

OV-102 = 180 total

Work Days launch




to repair (MTTR)


49

9/28/2010

Space Shuttle Launch Information Data Base Continued

Space Shuttle Launch Information Data Base

Mis

sio

n #

Orb

iter

102 &

M

issio

n F

lt. H

rs.

Orb

iter

102

Accum

. F

lights

Orb

iter

099 &

M

issio

n F

lt. H

rs.

Orb

iter

099

Accum

. F

lights

Orb

iter

103 &

M

issio

n F

lt. H

rs.

Orb

iter

103

Accum

. F

lights

Orb

iter

104 &

M

issio

n F

lt. H

rs.

Orb

iter

104

Accum

. F

lights

Orb

iter

105 &

M

issio

n F

lt. H

rs.

Orb

iter

105

Accum

. F

lights

Laun

ch P

ad #

Orb

iter

102 &

Pad

Work

day

Exposure

Orb

iter

102 &

OP

F

Work

day

Exposure

Orb

iter

102 &

VA

B

Work

day

Exposure

Orb

iter

099 &

Pad

Exposure

Hrs

.

Orb

iter

099 &

OP

F

Work

day

Exposure

Orb

iter

099 &

VA

B

Work

day

Exposure

Orb

iter

103 &

Pad

Exposure

Hrs

.

Orb

iter

103 &

OP

F

Work

day

Exposure

Orb

iter

103 &

VA

B

Work

day

Exposure

Orb

iter

104 &

Pad

Exposure

Hrs

.

Orb

iter

104 &

OP

F

Work

day

Exposure

Orb

iter

104 &

VA

B

Work

day

Exposure

Orb

iter

105 &

Pad

Exposure

Hrs

.

Orb

iter

105 &

OP

F

Work

day

Exposure

Orb

iter

105 &

VA

B

Work

day

Exposure

Sh

uttle

vehic

le

tota

l gro

und-f

low

w

ork

days.

STS 1 54.34806 1 LC 39A 104 Days 531 Days 33 Days 668 days





STS 6 122.2403 1 LC 39A 115 Days 123 Days 6 Days 244 Days




STS 41B 191.2653 4 LC 39A 22 Days 52 Days 6 Days 80 Days

STS 41C 167.6686 5 LC 39A 18 Days 31 Days 4 Days 53 Days

STS 41D 144.9344 1 LC 39A 72 Days 123 Days 15 Days 210 Days

STS 41G 197.3925 6 LC 39A 22 Days 53 Days 5 Days 80 Days

STS 51A 191.7489 2 LC 39A 17 Days 34 Days 5 Days 56 Days


STS 51D 167.9231 4 LC 39A 15 Days 53 Days 5 Days 73 Days


STS 51G 169.6478 5 LC 39A 14 Days 37 Days 7 Days 58 Days

STS 51F 190.7572 8 LC 39A 31 Days 39 Days 5 Days 75 Days

STS 51I 170.295 6 LC 39A 22 Days 27 Days 7 Days 56 Days

STS 51J 97.74389 1 LC 39A 34 Days 84 Days 14 Days 132 Days

STS 61A 168.7475 9 LC 39A 14 Days 35 Days 4 Days 53 Days



STS 51L 0.020278 10 LC 39B 28 Days 30 Days 5 Days 63 Days

STS 26 97.00306 7 LC 39B 88 Days 221 Days 13 Days 322 Days








STS 36 106.3061 6 35 Days 69 Days 6 Days 110 Days





50

9/28/2010


Mis

sio

n #

Orb

iter

102 &

M

issio

n F

lt. H

rs.

Orb

iter

102

Accum

. F

lights

Orb

iter

099 &

M

issio

n F

lt. H

rs.

Orb

iter

099

Accum

. F

lights

Orb

iter

103 &

M

issio

n F

lt. H

rs.

Orb

iter

103

Accum

. F

lights

Orb

iter

104 &

M

issio

n F

lt. H

rs.

Orb

iter

104

Accum

. F

lights

Orb

iter

105 &

M

issio

n F

lt. H

rs.

Orb

iter

105

Accum

. F

lights

Laun

ch P

ad #

Orb

iter

102 &

Pad

Work

day

Exposure

Orb

iter

102 &

OP

F

Work

day

Exposure

Orb

iter

102 &

VA

B

Work

day

Exposure

Orb

iter

099 &

Pad

Exposure

Hrs

.

Orb

iter

099 &

OP

F

Work

day

Exposure

Orb

iter

099 &

VA

B

Work

day

Exposure

Orb

iter

103 &

Pad

Exposure

Hrs

.

Orb

iter

103 &

OP

F

Work

day

Exposure

Orb

iter

103 &

VA

B

Work

day

Exposure

Orb

iter

104 &

Pad

Exposure

Hrs

.

Orb

iter

104 &

OP

F

Work

day

Exposure

Orb

iter

104 &

VA

B

Work

day

Exposure

Orb

iter

105 &

Pad

Exposure

Hrs

.

Orb

iter

105 &

OP

F

Work

day

Exposure

Orb

iter

105 &

VA

B

Work

day

Exposure

Sh

uttle

vehic

le

tota

l gro

und-f

low

w

ork

days.


STS 37 143.5456 8 LC 39B 22Days 97 Days 6 Days 125 Days








STS 49 213.2939 1 LC 39B 49 Days 217 Daysa 6 Days 272 Days
































STS 79 243.3244 17 LC 39A 25 Day 73 Days 17 Days 115 Days



51

9/28/2010


Mis

sio

n #

Orb

iter

102 &

M

issio

n F

lt. H

rs.

Orb

iter

102

Accum

. F

lights

Orb

iter

099 &

M

issio

n F

lt. H

rs.

Orb

iter

099

Accum

. F

lights

Orb

iter

103 &

M

issio

n F

lt. H

rs.

Orb

iter

103

Accum

. F

lights

Orb

iter

104 &

M

issio

n F

lt. H

rs.

Orb

iter

104

Accum

. F

lights

Orb

iter

105 &

M

issio

n F

lt. H

rs.

Orb

iter

105

Accum

. F

lights

Laun

ch P

ad #

Orb

iter

102 &

Pad

Work

day

Exposure

Orb

iter

102 &

OP

F

Work

day

Exposure

Orb

iter

102 &

VA

B

Work

day

Exposure

Orb

iter

099 &

Pad

Exposure

Hrs

.

Orb

iter

099 &

OP

F

Work

day

Exposure

Orb

iter

099 &

VA

B

Work

day

Exposure

Orb

iter

103 &

Pad

Exposure

Hrs

.

Orb

iter

103 &

OP

F

Work

day

Exposure

Orb

iter

103 &

VA

B

Work

day

Exposure

Orb

iter

104 &

Pad

Exposure

Hrs

.

Orb

iter

104 &

OP

F

Work

day

Exposure

Orb

iter

104 &

VA

B

Work

day

Exposure

Orb

iter

105 &

Pad

Exposure

Hrs

.

Orb

iter

105 &

OP

F

Work

day

Exposure

Orb

iter

105 &

VA

B

Work

day

Exposure

Sh

uttle

vehic

le

tota

l gro

und-f

low

w

ork

days.

































STS 107 382.35 28 LC 39A

Totals: 7217.821 28 1497.767 10 5805.566 30 5278.061 26 5122.486 19 1076 Days 2979 Days 164 Days 311 Days 454 Days 48 Days 964 Days 3019 Days 192 Days 838 Days 2488 Days 134 Days 611 Days 2169 Days 93 Days15,831

Days

Grand Total

Information

7,218Hrs.

with 28 flts

1,498Hrs.

with 10

flits

5,806Hrs.

with 30 flts

5,278Hrs.

with 26 flts

5,123Hrs.

with 19

flits

If 6 of 7

days/wk

are work

days, avg.

days @

pad/launch

= 44.8

If 6 of 7

days/wk

are work

days, avg.

days @

pad/launch

= 36.3

If 6 of 7

days/wk

are work

days, avg.

days @

pad/launch

= 37.5

If 6 of 7

days/wk

are work

days, avg.

days @

pad/launch

= 37.6

If 6 of 7

days/wk

are work

days, avg.

days @

pad/launch

= 37.5

Total continuous ground processing Work Days resulting in Launch-on-time (without scrub turnaround) for OV-105 was 1546 days in 5 flight

groupings or a mean-time-between-failure (MTBF) = 309.2 Work-days










52

9/28/2010

Appendix III

Space Shuttle Fluids Overview Matrix

CHEMICAL AND PHYSICAL CHARACTERISTICS

CONTROLLED BY REFERENCE DOCUMENT

SE-S-0073 Revision G (6-23 CHANGE NO. 77)

Main

Pro

puls

ion (

Orb

iter,

ET

and S

SM

E)

Aft P

ropuls

ion O

rbit

Maneuvering a

nd

Reaction C

ontr

ol

Reaction C

ontr

ol

Auxili

ary

Pow

er

Unit -

Orb

iter

(AP

U)

Auxili

ary

Pow

er

Unit -

SR

B (

AP

U)

SS

ME

and O

rbiter

Hydra

ulic

SR

B H

ydra

ulic

Fuel C

ells

/Cry

o

Environm

enta

l and

Therm

al C

ontr

ol / Life

Support

Purg

e, V

ent, a

nd D

rain

~ #

of uniq

ue s

ub-

syste

ms r

equirin

g s

erv

ice

each flig

ht

~ #

of in

div

idual in

terf

aces

requirin

g s

erv

ice (

connect

& d

isconnect)

for

each

flig

ht (n

ot in

clu

din

g

multip

le g

round s

tations)

Revised R

col values

DRAFT

SERVICING FluidRef. Table

SE-S-0073

Maximum

Sub-system

Cleanliness

Level Req'd

Operational Assembly

Cleaning

Verification &

Testing

Propellant Pressurizing Agent, Helium 6.3-1 X X 100 A X (100A) X (200) X (200) X (100) NA NA X (100) X (200A) X (300A) 17

Liquid Oxygen 6.3-2 X 800 A X (800A) 1

Gaseous Nitrogen 6.3-3 X X X 100 A X (100A) X (200) X (200) X (100) X (100) X (200) X (200) 6 20 20

Gaseous and Liquid Oxygen 6.3-4 X 200A X (200A) X (200A) 2 10 5

Gaseous and Liquid Nitrogen 6.3-5 X X 200A X (200A) X (200A) X (Class 5,000) 5

Gaseous and Liquid Hydrogen 6.3-6 X 200 X (400) X (200) 4 10 5

Hydraulic Fluid 6.3-7 X 190 X (190) X (190) 7 15 7

Water (Grade A or B) 6.3-8 X X 100/200A X (1000) X (200A) X (200A) X (100) X (100) X (200) X (200) X ( 200A) 3

Propellant, Monomethylhydrazine 6.3-9 X 200 X (200) X (200) 5

Propellant, Nitrogen Tetroxide (MON-3) 6.3-10 X 200A X (200A) X (200A) 5

Propellant, Hydrazine 6.3-11 X X 100 X (100) X (100) 7

Argon 6.3-12 X X 100A X (100A) X (200) X (200) X (100) NA X (200A) X (100A) X (Class 5,000) 2

Propellant, Mixed Oxides of Nitrogen (Deleted) 6.3-13 0

Lubricating Oil 6.3-14 X 300 X (300) X (300) 7

Conditioned Air - Purge, Vent, and Drain 6.3-15 X X Class 5,000 X (1000) X (Class 5,000) 5

Potable Water 6.3-16 X 300 X (300) 1

Carbon Dioxide 6.3-17 X X 100 NA X (100) 0

Ammonia 6.3-18 X 300A X (300A) 2

Biocide Flush Fluid 6.3-19 X 300 X (300) 1

Refrigerant 21 (New and Recycled) 6.3-20 X X 300 X (300) 3

Avionics Fire Extinguishing Fluid 6.3-21 X NA 2

Conditioned Air - ECS 6.3-22 X Class 100,000 X (Class 100,000) 5

Fuel Cell Cooling Fluid 6.3-23 X 300 X (300) 3

Breathing Oxygen 6.3-24 X 200A X (200A) 0

Isopropyl Alcohol 6.3-25 X 100/200A X (200) X (200) X (100) X (100) X (100) X (100) X (200A) X (200) 0

Trichlorotrifluoroethane 6.3-26 X 200A X (400A) X (200A) X (200A) X (200A) X (200A) X (200A) 0

Denatured Ethyl Alcohol 6.3-27 X 300 X (300) 0

Trichloroethylene 6.3-28 X 800A X (800A) 0

Breathing Air Mixture 6.3-29 X 200A X (200A) X (200A) 0

Airlock LCG Cooling Water 6.3-30 X 300 X (300) 2

Ferry Flight WCL Fluid 6.3-31 X 300 X (300) 0

Heat Transport Water 6.3-32 X X 200/300A X (200) NA X (300A) 3

Shock Strut Hydraulic Fluid - Orbiter 6.3-33 X 190 X (190) NA 0

Missile Grade Air (Deleted) 6.3-34 0

Refrigerant 114 6.3-35 X 300 X (300) 0

Breathing Air (EMU Ground Test Only) 6.3-36 X 200A X (200A) 0

Nitric Oxide (Deleted) 6.3-37 0

EMU Gaseous Oxygen 6.3-38 X 100A X (100A) 2

Propellant, Nitrogen Tetroxide (MON-10) 6.3-39 0

RTG Heat Transport Fluid 6.3-40 X 300 X (300) 0

APU/WSB Pre-Flush Lubricating Oil 6.3-41 X 300 X (300) NA 0

WMS Flush Fluid (1) 6.3-42 X 300 X (300) 0

Waste Tank Cleaning Fluid (1) 6.3-43 X 300 X (300) 0

Isopropyl Alcohol 6.3-44 X 400A X (400A) 0

WMS Servicing Fluid (1) 6.3-45 X 300 X (300) 0

HCFC-225 6.3-46 X X 200A X (400A) X (200A) X (200A) ( 200A) X (200A) 0

ZBA25 6.3-47 X 400 X (400) 0

Refrigerant 124 (HCFC-124) 6.3-48 X X 300 X (300) 2

Perfluorohexane (PF-5060) (Lear Romec Quick Disconnects Only) 6.3-49 X 300 X (300) 0

Vertrel MCA (Decafluoropentane [62%] and Trans-Dichloroethylene [38%] 6.3-50 X 200A X (400A) X (200A) ( 200A) X (200A) X (200A) 0

HFE-7100 (Methoxy-nonafluorobutane) 6.3-51 X 200A X (400A) X (200) ( 200A) X (200A) X (200A) 0

Vertrel XF (1,1,1,2,3,4,4,5,5,5-Decafluoropentane) 6.3-52 X 400A X (400A) 0

HCFC-225 G 6.3-53 X 200A X (400A) X (200A) ( 200A) X (200A) X (200A) 0

PGME/Water Azeotrope (1) 6.3-54 X 200 X (200) NA 0

Note: Reference source does not address applications in Sub-systems control. 102 total

Shuttle Sub-system Application |---------------------------- Shuttle Sub-system Identification --------------------------|

SHUTTLE VEHICLE SERVICING FLUIDS


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Appendix IV

Hardware Dependability Influence on Propellant Cost/LCC Goals

Propellant Cost Influence In Achieving

Advanced Reusable Space Transportation Systems Cost/Pound Goals

STATEMENT OF THE CHALLENGE

With the goal being $100. /# of payload to orbit, the first reactions for most is to assume the goal is

unrealistic; therefore, establish a range with a new must criteria of $500. /# to orbit. However, effectively

applying all of our engineering experience and expertise, the space community may be surprised at the

progress toward achieving this goal.

This challenging goal ($100. /# of payload to orbit) clearly suggests we must change ―the way we do

business‖ today. The program must establish the requirements to provide the controls required for

achievement, i.e., the offline maintenance function alone represents ~50% of this goal regardless whether

flying with normal boiling point or densified hydrogen and without considering the added cost of the

purge gas requirements. The maintenance function must be scaled back by requiring the propulsion

system fly considerably more missions between depot cycle and be more airline like the commercial gas

turbines. This will not only drastically reduce the propellant cost, but, at the same time will bring down

the logistics repair parts cost to be more inline with our goal of becoming more ―airline like‖ the flight

frequency will challenge us to provide propellant on site recycling our losses and at a much-reduced cost.

INTRODUCTION

Propellant cost is heavily driven by the way we are doing business, i.e., propulsion design philosophy.

Design characteristics such as mixture ratio, design life, hardware depot cycle frequency, and propellant

mass/ volume characteristics all influence a space transportation systems propellant cost. Also be aware

that the user pays for all the losses and the price is a function of quantity used because there are fixed and

variable cost associated with producing the product. Therefore, an example is provided using the Shuttle

program Liquid Hydrogen cost of calendar year 2001 for a reference case. SSC testing represents a Depot

cycle of the LH2 turbo-pump after 3 flights, the LO2 turbo-pump after 10 flights and 3 new SSME green-

runs. The SSME requires depot cycle after 20 flights. This insight should be useful in establishing the

Level I requirements for advanced space transportation systems.

STS LH2 USAGE FOR CALENDAR YEAR 2001

Liquid Hydrogen is used at MSFC, SSC, and KSC for the Shuttle program.

1. Quantity procured for MSFC was 406,406 #’s

2. Quantity procured for SSC was 5,793,771 #’s

3. Quantity procured for KSC was 2,213,738 #’s

4. Quantity actually flown by the Shuttle was 1,366,800 #’s in CY 2001 for 6 launches

Liquid Hydrogen purchased for KSC was by separate contract and the purchase price was $1.97/# with

61.74% of propellant launched. Remainder was transfer and storage losses. If this were the only added

cost, the LH2 would be $3.19/# launched. If the concept of operation were to assume that the hydrogen


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flown were at triple point in stead of normal boiling point, the realized propellant launched is estimated to

be only 40% that purchased which would drive the price to at least $4.93/# launched. However, this is

only a part of the actual cost of propellant (densified hydrogen) when assessing the program as a whole.

Liquid Hydrogen purchased for MSFC and SSC was by separate contract and the purchase price because

of the large quantity was ~ $1.09/# with NO propellant launched by Shuttle. This propellant was used for

hardware depot maintenance (each SSME was refurbished and tested after ever flight during the calendar

year 2001) and technology and reliability testing functions. Therefore, the maintainability burden

designed into the hardware has a direct influence on the propellant logistics cost to operate a Space

Transportation System, e.g., the greater time between depot maintenance cycles and hardware repair

actions used in determining their operability requirements will greatly reduce the program recurring cost.

To control and minimize the recurring cost, these operability requirements must be specified at Level I.

STS PAYLOAD COST PER POUND TO ORBIT FOR LH2 ONLY AND ACTUAL LH2 COST PER POUND LAUNCHED

When factoring in all uses of Liquid Hydrogen for the Shuttle propulsion program, it was determined that

8,413,915 #’s of Liquid Hydrogen were purchased and only 1,366,800 #’s flown on Shuttle in calendar

year 2001. Therefore, only 16.24% of the propellant purchased was launched. This causes the actual

Liquid Hydrogen cost to be $8.13/# launched. The propellant launched would have been 61.74% with NO

maintenance functions required for the calendar year 2001 with 6 flights flown. Liquid Hydrogen cost

would have been $3.19/# launched or $14.54/ # of payload to orbit assuming a 50,000 # payload. When factoring in the maintenance cost of the propulsion systems the cost is $8.13/ # launched and

$23.86/ # of a 50,000 # payload to orbit.

A NOTIONAL SCENARIO FOR 3rd

GEN REQUIREMENTS CONSIDERATION

To project a scenario for 3rd

Gen RLV with a SSTO vehicle and 20,000 #’s payload the actual cost

allocation would require $3.19 to $9.87/# for Liquid Hydrogen alone. Assume the vehicle was 1,113,636

#’s GLOW and was SSTO with a mass fraction of .88. This vehicle would require 140,000 #’s of Liquid

Hydrogen and 840,000 #’s of Liquid Oxygen assuming a 6:1 mixture ratio propulsion system design.

LIQUID HYDROGEN CONSIDERATIONS:

Considering the losses for storage and transfer to load the vehicle with normal boiling point hydrogen

the cost of Liquid Hydrogen would be $3.19/# launched or $22.33/# of payload to orbit at 20,000#’s

per flight, but if it is assumed that the vehicle requires the hydrogen loaded at the triple point conditions,

the added equivalent losses would drive the cost of Densified Liquid Hydrogen to $4.93/# launched or

$34.47/# of payload to orbit. These values do not include propulsion maintenance cost that will be

required if not corrected by leveling a requirement at Level I to control recurring cost of the program.

Therefore, if the level of hydrogen use for off-site engine depot maintenance remains the same, this 3rd

Gen RLV will cost $4.94/ # actual launched cost and would represent $34.60/# of payload to orbit

for the 20,000 # payload just to cover this maintenance function. Assuming this depot maintenance

function is not changed, this cost must be added to the launch site cost. Therefore, assuming the 3rd

Gen

RLV used normal boiling point hydrogen, we will add the maintenance cost of $4.94/# of hydrogen

to the launch cost of $3.19/# of hydrogen equaling $8.13/# of hydrogen launched or $56.93/# of

payload to orbit assuming a 20,000 # payload.


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COST INSIGHT OF TECHNOLOGY CONCEPT ALTERNATIVE

However, concepts requiring the propellant be loaded in a densified state near the triple point instead of

normal boiling point hydrogen will require more equivalent hydrogen at the launch site. This will drive

the Densified Liquid Hydrogen cost to $4.93/# launched or $34.47/# of payload to orbit for the

launch site only. Again if I assume the off site SSC + MSFC use were to remain the same and only using

normal boiling point hydrogen but flying the vehicle at these new triple point conditions, the added

required quantity at KSC would increase from ~226,757 #’s to ~350,000 #’s equivalent resulting in

the cost being $9.87/# actual launched cost and would represent $69.09/# of payload to orbit for the

20,000 #payload.

The above thinking does not include many factors that have not been determined. The concept of loading

the flight tank with densified liquid hydrogen may require the removal of its sensible heat using cold

hydrogen gas to avoid the large heat rise in the liquid during the loading. Also the flight tank will require

pressure stabilization with cold helium gas during the loading and replenish operations to avoid tank

collapse, as the Liquid Hydrogen will try to seek its equilibrium pressure of 1 psia. These functions were

accomplished by using LH2 cold bath heat exchangers that will add another large quantity of Liquid

Hydrogen at the launch site to accomplish this concept of operation.

LIQUID OXYGEN CONSIDERATIONS:

If a similar loss ratio is assumed for Liquid Oxygen, now determine the cost associated with the oxygen

propellant for this 3rd

Gen RLV scenario.

It was stated earlier that the 1,113,636 #’s GLOW was a SSTO with mass fraction of .88 with the mixture

ratio of 6:1 and a payload of 20,000 #’s to orbit. Therefore, the Liquid Oxygen load would be 840,000

#’s. With the 61.74% propellant launched value used for Liquid Hydrogen being considered the same the

Liquid Oxygen procured would be 1,360,544 #’s and at a price of $132.80/ton the cost would be

$90,340.14 or $0.11/# launched - $4.52/# of payload launched for the launch site cost.

Again factoring in the propulsion maintenance cost would increase this cost.

The propellant purchased off site for this maintenance and R&D function is 2.8 time that at KSC for

Shuttle. Therefore, using this ratio to determine the quantity purchased off site would be 3,809,524 #’s of

Liquid Oxygen. Assuming the same price as KSC procurement would result in cost of $252,952.38 or

$0.30/# launched - $12.65/# of payload to orbit for this maintenance and R&D function.

Therefore, combine these two functions to determine the cost of continuing business as usual would result

in the following cost of $0.11/# launched at KSC plus the $0.30/# launched for maintenance would equal

$0.41/# launched or $17.22/# of payload to orbit for Liquid Oxygen.


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LIQUID OXYGEN AND HYDROGEN COMBINED CONSIDERATIONS:

Now combine the cost of both Liquid Hydrogen and Oxygen for this 3rd

Gen RLV scenario:

1. Assume the use of normal boiling point propellant only and only that required at the launch site.

LH2 @ $3.19/# launched or $22.33/# of payload to orbit plus LO2 @ $0.11/# launched or $4.52/#

of payload to orbit = total of $26.85/# of payload to orbit as launch site cost only.

2. Consider the combined cost for the propulsion maintenance and R&D function off site. LH2 @

$4.94/# launched or $34.60/# of payload to orbit plus LO2 @ $0.30/# launched or $12.65/# of

payload to orbit = total of $47.25/# of payload to orbit considering only normal boiling point

propellant use off site for the maintenance and R&D function.

3. Now combine the launch site and off site maintenance functions to determine the actual cost of

doing business as with the Shuttle. The LH2 cost will now be $3.19/# + $4.94/# = $8.13/#

launched or $56.93/# of payload to orbit. The LO2 cost will be $0.11/# + $0.30/# = $0.41/#

launched or $17.17/# payload to orbit. When combining the two propellants we have $8.13/# +

$0.41/# = $8.54/# of propellant launched or $56.93/# + $17.17/# = $74.10/# of payload to orbit

cost using normal boiling point propellant.

4. Assuming the use of densified LH2 only at the launch site, but business as usual for the

maintenance function, the LH2 cost will now be $4.93/# + $4.94/# = $9.87/# launched or $69.07/#

of payload to orbit. The LO2 cost will be $0.11/# + $0.30/# = $0.41/# launched or $17.17/#

payload to orbit. When combining the two propellants, we have $9.87/# + $0.41/# = $10.28/# of

propellant launched or $69.07/# + $17.17/# = $86.24/# of payload to orbit. This resulting cost is

using densified hydrogen at the launch site and normal boiling point propellant off site for

maintenance functions.

The desire to use densified liquid oxygen would result in a similar cost increase relationship for the

LO2 propellant.

OTHER CONSIDERATIONS NOT INCLUDED IN THE ABOVE ASSESSMENT:

The above propellant cost assessment only represent ~80% of the propellant cost when working with

cryogenics as we do business today. As an example, a data sample is provided to support this conclusion

from SSC SSME operations during the FY 2000 period.

Liquid hydrogen cost = $7,314,922. ------- 70.67% of total propellant cost

Liquid oxygen cost = $1,323,150 ---------- 12.78% of total propellant cost

Gaseous Helium cost = $1,044,120 -------- 10.09% of total propellant cost

Liquid nitrogen cost = $668,230 ------------- 6.46% of total propellant cost

Liquid propellants = 83.45% and purge requirements represent 16.55% of total annual cost to support the

SSME maintenance and R&D operation off site during FY 2000. Liquid hydrogen use during FY-2000

was 19.5% greater than CY-2001: therefore, the above cost is higher than CY-2001 but cost relationship

should be considered the same. These purge gas cost will also be required at the launch site in similar

ratios as SSC operation for the SSME operation.


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CONCLUSIONS

To achieve the cost goal ($100. /# of payload to orbit), this analysis concludes the space community must

change the way we are doing business today to be more inline with airline operations. The design life,

time between depot cycles, and the hardware inherent reliability must all be improved and follow the

leadership of the airline/aircraft industry. Suggest increasing the depot maintenance cycle time to be more

airline like the commercial gas turbines, use combined cycle which induces air to replace some of the

oxygen, consider moving to oxygen rich combustion for the early phase of flight, and drive to single stage

to orbit, the propellant cost will be more in line with our objective goal. These improvements are required

just to reduce the fluid commodity cost required to achieve the overall cost objective. Also the propellant

cost of $74.10/# of payload to orbit does not include the gases required to support the cryogenic operation

and this was 16.5% of the cost to do business at SSC. The desire of some to use densified propellants

could increase the propellant cost to $86.24/# of payload to orbit plus a considerably higher cost for the

gases. To safely use densified propellants will require a cold helium bottle submerged in the propellant

tanks with the bubbling of helium to maintain the tank structure stability during ground servicing

operations. Again this choice is directly opposite the desired cost reduction from today’s way of doing

business.

13.97

22.33

2.97 4.52

14.7

41.55

56.62

98.17

0

10

20

30

40

50

60

70

80

90

100

Co

st/

# o

f p

ay

loa

d t

o o

rbit

Propellant Cost/# of payload to orbit for 20,000 #/flight

Notional Scenario

LH2 procurement cost $13.79 /# of payload LH2 procurement cost plus losses $22.33 /# of payload LO2 procurement cost $2.79 /# of payload LO2 procurement cost plus losses $4.52 /# of payload Helium & nitrogen cost for propulsion engine support support

$14.70 /# of payload Major propellant and gases cost with losses $41.55 /# of payload Propulsion Engine maintenance propellant & gases cost $56.62 /# of payload Total: Major propellant and gas and engine maintenance cost

$98.17 /# of payload


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Propellant Cost/# of payload to orbit for 20,000

#/flight Notional Scenario:

$13.79

$22.33

$2.79$4.520

$14.70 $16.58

$41.55

0

$56.62

$16.58

$98.17

$0.00

$20.00

$40.00

$60.00

$80.00

$100.00

$120.00

STS Maintenance frequency & methods vs.

reductions required to achieve Major Goals

Co

st/

# o

f p

aylo

ad

to

orb

it

O Total: Major propellant and gas and engine maintenance cost

Major propellant and gases cost with losses

LH2 procurement cost plus losses

x Helium & nitrogen cost for propulsion engine support

LO2 procurement cost plus losses *

Propulsion Engine maintenance propellant & gases cost

--Goal--

Propellant Use Reduction Fraction

Shuttle

STS

0

0 1

0 1


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Appendix V

Shuttle Reference Case 1: safety driven functional requirements

Number of safety driven functional requirements to maintain safe control of systems during

flight and ground operations:

o ET element:

Nose cone inerting heated GN2 purge

Intertank inerting heated GN2 purge

GH2 ground umbilical plate He purge

Hazardous gas detection system in the intertank

Hazardous gas detection (GH2 sensors) for GH2 ground umbilical plate

interface

Lox anti-geysering He bubbling system

o Orbiter – MPS

Lox POGO suppression system

Aft compartment GN2 purge

Aft compartment hazardous gas detection system

Orbiter/ET Lox umbilical He purge

Orbiter/ET LH2 umbilical He purge

Orbiter/ET LH2 umbilical hazardous gas detection system (GH2 sensors)

Orbiter/Ground Lox umbilical He purge

Orbiter/Ground LH2 umbilical He purge

Orbiter/Ground umbilical hazardous gas detection system (GH2 sensors)

LH2 main feedline manifold high point bleed system

o SSMEs

LH2 turbopump thermal conditioning system (3)

Lox turbopump thermal conditioning bleed system (3)

Lox turbopump seal He purge system (3)

GH2 lead flow burn-off ignition system (6)

SSME/MLP exhaust sound suppression system

MLP deck sound suppression system from SSME driven drift at liftoff

o Orbiter OMS/RCS

FRCS compartment GN2 purge

APS right side pod compartment GN2 purge

APS left side pod compartment GN2 purge

FRCS fuel umbilical system purge

FRCS ox umbilical system purge

APS right side fuel system umbilical purge for OMS

APS right side ox system umbilical purge for OMS

APS right side fuel system umbilical purge for RCS

APS right side ox system umbilical purge for RCS

APS left side fuel system umbilical purge for OMS

APS left side ox system umbilical purge for OMS

APS left side fuel system umbilical purge for RCS

APS left side ox system umbilical purge for RCS


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Operational personnel SCAPE suit system (plus maintenance systems for

support)

Toxic vapor detection system (many sensors & personnel badges @ 3

stations)

Hazardous waste management systems (used @ 3 major stations–OPF,HMF,

& Pad

o Orbiter PRSD

LH2 & GH2 umbilical plate He purge system (several and dependent of # of

tanks)

Lox & Gox umbilical plate purge system (several and dependent of # of tanks)

o SRB’s

Aft skirt GN2 purge (2)

Field joint heater system (10)

Ignition overpressure suppression & control system (one foot H2O coverage

of top of exhaust opening) (2)

Ignition overpressure suppression & control H2O injection of MLP exhaust

(2)


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Appendix VI

Shuttle Reference Case 2 for safety driven limited access control

Number of safety driven limited access control operations:

o Orbiter aft compartment

o ET Intertank

o SRB aft skirt (2)

o Handling SRB segments (8)

o Lifting and handling the Orbiter

o Lifting and handling the ET

o Installing and connecting ordnance in system (~200 in orbiter)

o Installing and connecting ordnance in SRB (~12)

o Installing/mating booster separation motors on Shuttle

o Installing and connecting ordnance on separation motors

o Servicing APU’s with hydrazine on Orbiter (3)

o Servicing APU’s with hydrazine on SRB’s (4)

o Propellant servicing OMS & RCS on Orbiter (3 locations @ PAD, 3 locations @

OPF, & HMF)

o Performing any maintenance on Orbiter OMS, RCS & APU’s at OPF & HMF (3

stations)

o Performing any maintenance on hypergolic systems and re-supply of propellants at

PAD (7)

o Performing recovery & recycle on SRB APU’s (2 stations)

o Servicing Orbiter NH3 & Freon 21 systems (5)

o Loading cryogenic propellant on the integrated Shuttle Orbiter/ET @ the pad (2)

o Replenishing the cryogenic (LH2 & Lox) propellant at the Pad storage tanks (4)

o Loading/servicing the Orbiter Fuel cell/PRSD cryogenic system at the PAD (Lox &

LH2)

o Preparing high purity Lox for the Orbiter fuel cell/PRSD system


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Appendix VII

Space Shuttle Shortfalls Assessment Results

The SPST Functional Requirements Subteam prepared the following table which shows the current capabilities of the Space Shuttle and the Critical

Shortfalls relative to the initial space shuttle requirements. It is the SPST position that NASA must understand the shortfalls of the current Space

Shuttle before NASA can correct these shortfalls or design them out of the next generations of space launch vehicles.

Current RLV Space Transportation Systems Shortfalls Assessment

Space Shuttle Critical Shortfalls Relative to Requirements

Attribute I D.


Desired Attributes needed to

Correct Shortfalls of

Current Systems


Critical Shortfalls


D 2.1 1. Program STS Design Life 10 years or

100 flts/veh.

20+ years, but only

30 flts/veh. Max., but still

counting

The Space Shuttle was intended to fly 10 flights

per year each without extensive maintenance and

recertification between flights (160 hour

turnaround). Design complexity and hardware

dependability only permitted less than 3 flights

per year. Avg. 100 components replacement plus

~ 400 expendable or limited life parts. Application

must be well understood so that the reliability

requirements flow-down supports the design life

after balancing the requirement with safety and

maintainability.


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Attribute I D.




Current Systems


Critical Shortfalls


D 2.1 2. Propulsion Main Engine Life 55 starts/Depot

cycle.

After 20+ years ops:

SSME depot cycle 20,

LO2 turbo-pumps 10 &

fuel turbo-pumps 3 flts

The SSME initial design life was 55

flights before entering depot cycle, but limited

life/dependable hardware has required extensive

labor, time, and engine depot support, e.g.,

Resulting in high cost per flight. Application must

be well understood so that the reliability

requirements flow-down supports the design life

after balancing the requirement with safety and

maintainability. Also the reliability requirement

must be demonstrated by testing and improved

until the requirement is met.

A 1.2 3. Recurring cost: $100.00/lb to LEO

@ 28.8 degrees

and 100 nmi

~ $10,000.00/lb to LEO

Actual Shuttle recurring

cost over the total 21

operating years =

~ $57.876 Billion.

Over the 21 operating

years 2,934,200 lbs has

been delivered to various

orbits. Ref. Appendix IX

The initial design recurring costs were

$100.00/lb to LEO which was based on achieving

an allocated 40 launches per year using 4 orbiters

flying 65,000 lbs per flight. Most flights do not fly

at maximum capacity and the 10 flights per year

per each orbiter was not achievable because of

complexity (optimizing at the sub-system level

―stove-piping‖ and not at the overall STS systems

level) and poor dependability of total system.


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Attribute I D.




Current Systems


Critical Shortfalls


A 1.1/

P 5.2

& 5.4

4. Non-recurring cost:(DDT&E

and Acquisition)

$5.0 Billion $15.0 Billion The initial design non-recurring cost

estimate were $5.0B based on an allocated

DDT&E schedule. Due to non-mature major

technologies (HP LH2/LO2 staged combustion

rocket engine, re-entry TPS, Solid rocket flex

nozzle seal, Ice/frost less cryogenic tanks, and

100% digital flight/ground control systems),

schedule was overrun 2 years because much

unplanned technology maturation was required.

R 3.2 5. Each vehicle flight rate: 10 flights/yr. 2.5-3 flights/yr. The initial design allocation for turnaround was



but because design requires the functional

integrity to be broken each flight to perform the

turnaround, ~ 400 expendable or limited life parts

to be replaced, and ~ 100 failed components to be

replaced during the turnaround operation, and

extensive servicing require (too many different

fluids and too many interfaces along with the

extensive support infrastructure), the achievable

flight rate is ~ 2.5 per year.


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Attribute I D.




Current Systems


Critical Shortfalls


R 3.2 6. Fleet flight rate: 40 flights/yr. at

ETR

10 flights/yr. The initial design allocation for turnaround was





~ 400 expendable or limited life parts to be

replaced, and ~ 100 failed components to be



fluids and too many interfaces with the extensive

support infrastructure), the achievable flight rate

is ~ 2.5 per year.


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Attribute I D.




Current Systems


Critical Shortfalls


R 3.1

&3.9

7. Vehicle turnaround time:

160 hours 1296 hours Min. The initial design allocation for turnaround was





~ 400 expendable or limited life parts to be

replaced, and ~ 100 failed components to be



fluids and too many interfaces along with the

extensive support infrastructure), the achievable

flight rate is ~ 2.5 per year. Because the integrity

of systems are compromised to provide for parts

change-out and to support turnaround operations,

the STS must be recertified for each flight.

R 3.1

&3.9

8. Vehicle/System terminal

countdown:

2 hours 8 hours plus The initial design terminal countdown was 2 hours

including main propellant loading, crew ingress

and egress checks. The initial design requirement

was 24 hours from standby status (VAB/T-0)

which included this 2 hour terminal countdown.

This requirement was deferred because of the

added DDT&E cost to provide the automation for

crew ingress/egress and MPS Lox transfer

capacity needed.


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Current Systems


Critical Shortfalls


R 3.2 9. System performance to LEO 65,000 lbs at ETR

@ 28.8 degrees

and 100 nmi

55,000 lbs at ETR @ 28.8

degrees and 100 nmi

This shortfall was a result of lack of sufficient

margins in performance of each variable, e.g.,

orbiter over weight, SSME Isp low, and the drive

to keep the ET production cost low. Example of

cost to remove ET weight was an additional

$20,000,000./ET unit for a 6,000 pound weight

reduction.

R 3.1/

S 4.2

10. Space flight rescue time:

24 hr. VAB/T-0 Not capable, but now

being considered again

since Columbia event.

Initial ground system design includes the

capability to change out the payload at the pad,

but 2 hour terminal countdown capability was not

implemented and the lack of meeting the fleet

flight rate caused a deferral of this requirement.

11. STS Dependability/Safety Not stated Program has lost 14 flight

crew and 2 ground

members and two orbiters.

The initial design reliability was 0.98 for 100

missions of each orbiter. No requirements were

established for loss of flight or ground crew

members and the impact of insufficient component

reliability and environment interaction was not

sufficiently considered and understood.

D 2.2/

S 4.1

11.1 Loss of Vehicle

0.98 0.964 for the Orbiter

0.962 for the SRB

Program did not consider cost impact of vehicle

Loss accompanied with down time for the

investigation and corrective action required for

re-flight. Importance of loss of vehicle and the

resultant impact on the program was not

considered with proper risk reduction actions.

Target metric value was deficient in determining

its overall impact on program.


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Current Systems


Critical Shortfalls


D 2.1/

S 4.1

11.2 Flight system program

reliability:

0.98 for 100

missions of each

orbiter or 500

missions total for

fleet of 5 orbiters.

0.964 for the Orbiter

0.962 for the SRB

The initial design was for the reliability of each

orbiter to be 0.98 for 100 missions. Program did

not consider cost impact of vehicle loss

accompanied with down time for the investigation

and corrective action required for re-flight.

Importance of loss of vehicle and the resultant

impact on the program were not considered with

proper risk reduction actions. Target metric value

was also deficient in determining its overall

impact on the maintainability burden (plus large

depot maintenance and supply chain support)

resulting in reduced flight rate and increased cost

per flight.

D 2.3 11.3 Mission reliability:

The subject in

general was

controlled by NHB

5300.4

~0.96 All flight vehicle subsystems (except primary

structure, thermal protection system, and pressure

vessels) shall be established on an individual

Sub-systems basis, but shall not be less than

fail-safe. Safety, reliability, and maintainability

were controlled separately by NHB 5300.4 (lD-1),

August 1974.


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Current Systems


Critical Shortfalls


R 3.7 11.4 Maintainability:

Shuttle orbiter was

designed for 100

flights or 10 years

without planned

maintenance. No

other direct

requirement other

than (160 hours

turnaround)

except the Shuttle

SSME 55

starts/Depot cycle.

SRB was to be

recovered and

refurbished every

flight.

Replace Avg. of 100

components/flight

unplanned and best case

orbiter turnaround is ~ 960

hrs. There are many

limited life components on

the Shuttle orbiter, e.g., ~

200 expendable ordinance

items and

~ 200 other limited life

items to track & replace.

Also after 20+ years of

ops, SSME depot cycle is

every 20 flights, with the

LO2 turbo-pumps after

every 10 and fuel turbo-

pumps after every 3 flights

Only requirement was the 160 hour turnaround

and maintainability design efforts were dropped

early in the DDT&E phase because of cost

overruns and schedule concerns. Critical

component redundancy was implemented with

component reliability levels that ignored the

resultant maintainability burden. This lack of

controlled maintainability requirements, has

balanced with safety and reliability, contributed to

the large resultant cost per flight and the low flight

rate. No specific maintainability requirements set.

R 3.7 11.5 Component replacement

time or mean time to

repair (MTTR)

No Requirement

documented other

than 160 hours

turnaround

Example of SSME

Controller replacement

during scrub-turnaround:

Up to 5 days or 80 Hrs.

Only requirement was the 160 hour turnaround

and maintainability design efforts were dropped

early in the DDT&E phase because of cost

overruns and schedule concerns. This lack of

controlled maintainability requirements

(accessibility, intrusive nature of most of the

hardware and no automated functional verification)

has contributed to the large resultant cost per

flight and the low flight rate.


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Current Systems


Critical Shortfalls


R 3.1

&3.9

11.6 Launch Availability:

24 hour notice to

launch from

standby status

(VAB/T-0) VAB

rollout including

payload change-out

at the pad, MPS

propellant loading,

crew ingress, final

close-out checks

and terminal count.

14 Work Days at the Pad

is best case before STS-

51L and 19 Work Days at

the Pad has been

demonstrated after the

STS-51L event.

Requirements flow-down were not developed,

implemented and controlled to provide this

capability. Lack of major system integration

resulted into too many flight/ground service

interfaces, controlled access conditions and

extensive time consuming operations.

S 4.1

&4.2

11.7 Flight abort during ascent:

Prevent Loss of Crew

Designed for:

return-to-launch-

site; abort-to-orbit;

abort-once-around

(RTLS/ATO/AOA)

Did not demonstrate

RTLS or transatlantic

landings (TAL’s) and

ATO was required only

once, but did not result in

an aborted operation.

Abort during ascent operations required the

SRB’s to burn to completion and failure occurred

with the SRB resulting in the loss of the orbiter

(099) and its crew. Therefore, abort during

ascent did not cover all critical failure modes.

S 4.1

&4.2

11.8 Flight abort from orbit:

Prevent Loss of Crew

KSC prime, EAFB

secondary, &

several

contingencies

3 landing sites used:

(KSC, EAFB, and White

Sands)

No abort was provided during the descent phase

and an orbiter (102) and its crew were lost during

re-entry.


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Current Systems


Critical Shortfalls


D 2.4/

R 3.1

12. Flight environment: Launch

and Landing

95 percentile

natural

environment

expected at

operational

locations

7 mile visibility and no

rain

Orbiter TPS cannot function in design

environment without damage. TPS needs to be

more robust to be in compliance with

requirements and to avoid launch and landing

scrub/delayed operations.

R 3.1

&3.8

13. Mission Planning Cycle

Was considered

within the 40

flights/yr. with 4

vehicle fleet and

the 24 hr. notice to

launch

requirement.

400 day typical cycle Standardized payload accommodations were not


mechanical and fluids accommodations are

custom designed for every mission. Also

standardized mission planning for payload mass,

orbit destinations, etc. was not provided;

therefore, each mission is planned as a custom

mission.


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Current Systems


Critical Shortfalls


R 3.9/

A 1.5

14. Total number of assembly

functions required at the

launch site between flights

The shuttle initial

design requirement

provided an

allocation of 34

hours of the 160

turnaround for the

space vehicle

assembly. Also a


flight rate with 4

vehicle fleet size.

The two SRB stages are

completely assembled

from scratch at the launch

site for each launch on the

MLP, a new ET is

received and integrated

into the SRB/MLP stack,

with the Orbiter being

integrated as the final step

of building the flight

vehicle. The Orbiter

requires re-configuration

for each unique payload

structural attachment as

well as providing unique

airborne support

equipment to service the

payload after installation

into the Space Vehicle.

The large SRB vehicle element concept does not

lend to the objectives of an RLV that achieves a

40 launch per year flight rate as it must be built-

up at the launch site and the recovery operations

are more like salvage and reconstruction

operations. Design concept choice was

inappropriate for the objective of the space

transportation system.

Lack of performance margin drives a custom

redesign for payloads for each mission.


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Current Systems


Critical Shortfalls


R 3.9/

A 1.5

15. Total number of expendable

items/components included

in the reusable system design

Objective was to

provide an RLV

with a 160 hours

turnaround

capability that

could fly 40 time a

year with a fleet of

4 orbiters.

~ 200 ordinance items

replaced every flight and ~

200 other one-flight

limited life items on the

orbiter plus the

expendable ET and much

expendable hardware on

the SRB’s.

Designers and program managers loss sight of

the objective to build an RLV because of an

overriding focus on meeting the performance

requirements in the absence of any other

structured engineering management processes

used to control such thing as life cycle cost or

any of the other program level one objectives.

R 3.9/

A 1.5/

R 3.6

16. Total number of sub-systems

requiring servicing with

dedicated ground systems,

e.g., number of dif. Fluids,

number of dif. Electrical

supplies, etc.

Objective was to

provide an RLV

with a 160 hours

turnaround

capability that

could fly 40 time a

year with a fleet of

4 orbiters and with

a 24 hr. notice to

launch capability to

accommodate

rescue.

Shuttle System requires

the tracking and managing

of ~ 54 different fluids and

~ 30 unique fluids are

serviced every flight.

Many of these fluids are

common from one

discipline to another,




vehicle.

Not considered was a need to provide any control

as would have been considered over

constraining. Designers and program managers

loss sight of the objective to build an RLV

because of an overriding focus on meeting the

performance requirements in the absence of any

other structured engineering management

processes used to control such thing as life cycle

cost or any of the other program level one

objectives. Also the STS was optimized at the

sub-system level and not at the overall integrated

level. Electrical functions are custom managed

on the ground and uniquely provided through

separate umbilicals instead of simplifying the

flight to ground interface functions by providing

the electrical management on the vehicle.


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Current Systems


Critical Shortfalls


R 3.9/

A 1.5/

R 3.6

16. Total number of sub-systems

requiring servicing with

dedicated ground systems,

e.g., number of dif. Fluids,

number of dif. Electrical

supplies, etc.

Con’t

The Shuttle has ten (10)

major sub-system













flight. Data bus and


well as unique

instrumentation that have

not been accounted for in

this assessment.


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Current Systems


Critical Shortfalls


R 3.9/

A 1.5



mission

The shuttle initial

design requirement

provided an

allocation of 96

hours of the 160

turnaround for the

orbiter turnaround

including the

payload installation and

verification. Also a


flight rate with 4

vehicle fleet size.

Each different payload

requires the Orbiter to be

custom built to support the

structural load and any

servicing requires special

airborne support

equipment to be installed

and verified along with

optimizing the mass

impact on the payload for

these services. Also flight

software must be custom

built for each mission.

Standardized payload accommodations were not


mechanical and fluids accommodations are

custom designed for every mission. Also

standardized mission planning for payload mass,

orbit destinations, etc. were not provided;

therefore, each mission is planned as a custom

mission.


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Current Systems


Critical Shortfalls


R 3.9/

A 1.5

18. Total number of manual



critical flight functions, e.g.,

CG, fluid residuals content

and purity, functionality of

primary and backup system

hardware

No Requirement

documented other

than (160 hours

turnaround, 24 hr.

notice to launch,

and 40 flights/yr.

with 4 vehicle

fleet). Not

considered as a

need to provide

any control and

would be

considered as over

constraining.

There were no structured engineering

management processes put in place to provide

constraints or to limit these functional

requirements for each flight. There was no

automated functional verification capability

(IVHM) provided to reduce the labor

intensiveness of the task.

D 2.5 19. Launch on time (No launch

scrubs)

No Requirement

documented except

24 hr. notice to

launch for space

rescue and military

needs.

65 of the 113 missions

launched the day

scheduled (57.5%). Of the

48 launch scrubs, 13 were

weather related (27%)

However, some missions

were scrubbed more than

once/mission.

There was no requirement against reliability to

accomplish either the launch on time or meet the

24 hour notice to launch for a space rescue. Not

considered as a need to provide any control and

was considered as over constraining.


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Current Systems


Critical Shortfalls


D 2.6 20. Materials, fluids, and designs

properties and limitations

well understood through

failure with narrow

tolerances

Not considered as a

need to provide

any control and

would be

considered as over

constraining and

would have drove

up the DDT&E

cost considerably.

Because the limits are not

known, operational

controls must provide

margins to avoid

unplanned events.

Performance also carries

an extra margin to allow

for these uncertainties.

Program objectives were compromised because

of these added limitations do to uncertainties.

A 1.7/

P 5.2

21. LCC must be well defined

and understood by analysis

without any

allocations/assumptions so

that the business case closes

No Requirement

documented, but

the targeted NRC

was $5 billion and

the Recurring

target was $6.5

million/flight = to

$2.6 billion in 10

yrs. Or a total LCC

= to $7.6 billion.

Assumptions were

65,000 lbs to orbit

each flight, 10

flights/orbiter or 40

flights/yr at

$100/lb to orbit.

Shuttle NRC (DDT&E) =

$15 billion and the RCC

avg is ~$2.756 billion/yr.

Therefore, the intended 10

year program LCC would

have = $42.56 billion. But

the Actual Shuttle

recurring cost over the

total 21operating years = ~

$57.876 Billion. Or a total

LCC = $72.876 billion.

These actual costs do not

include any R&T cost

prior to the STS ATP (1-5-

72), e.g., SSME, TPS, etc.

A very large shortfall exist in the LCC

projections because they were based on

allocations that never came into fruition, e.g.,

65,000lbs to orbit each flight and 40 launches per

year using 4 orbiters. Also the DDT&E cost

projection had a large shortfall because of the

immature technologies causing an extended

schedule for this activity. Allocations of the

operational functions could not be met because

there was no engineering management processes

put in place to provide the necessary control

required.


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Critical Shortfalls


P 5.3 22. All technologies must be

matured at the TRL-6 level

or above and options must be

available as backups where

risk is moderate or above

prior to start of acquisition.

No Requirement

documented.

DDT&E schedule

and cost risks were

not considered a

necessity as we

were still working

to the Apollo

paradigm.

Five Major System’s

Technologies less than

TRL-6 level at ATP: High

Pressure LO2/LH2 Staged

Combustion Rocket

Engine, Vehicle TPS,

Large Solid Rocket Motor

Nozzle Flex Seal TVC

system, Ice/frostless

cryogenic tanks, & 100%

digital control

flight/ground systems

NASA started a major program with five major

non-mature technologies and no technology

backups. Also did not provide the margin in cost

or schedule to account for this high risk

approach.

S 4.3 23. All processes and operations


environmental regulations

and laws

No Requirement

documented?

Today’s

regulations were

not established

during the Shuttle

concept and

DDT&E phases

and it was assumed

NASA would abide

by the country’s

laws.

Stringent Environmental

/OSHA requirements have

been imposed since

Shuttle ATP

STS Program didn’t include cost allowances for

changes in the environmental laws.


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Critical Shortfalls


R 3.6 24. Vehicle, payload, and

ground systems integration

functions must be

compatible with all LCC


without any assumptions.

No Requirement

documented other

than (160 hours

turnaround, 24 hr.

notice to launch,

and 40 flights/yr.

with 4 vehicle

fleet). Not

considered as a

need to provide

any control and

would be

considered as over

constraining.


402 functional interfaces

of which the Propulsion

discipline alone has 236 of

which the SSME has




system for the - single


station (ICD-2-1A002) –

for the vehicle to ground

design and operations

activities. Note: The

orbiter element has ten

(10) more facility station

ICD’s at the launch site.

The above is an example

of all major flight element

interface support

requirements as the SRB’s

have 16 safety driven

functional requirements

and 28 safety driven

limited access control

requirements.





realistic and based on the functional

requirements along with assurance the objectives

would be met.


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Current Systems


Critical Shortfalls


R 3.6/

R 3.9

25. Easily Supportable

(minimum support

infrastructure required) and

must be compatible with all

LCC requirements by


assumptions

No Requirement

documented other

than (160 hours

turnaround, 24 hr.

notice to launch,

and 40 flights/yr.

with 4 vehicle

fleet). Not

considered as a

need to provide

any control and

would be

considered as over

constraining.


402 functional interfaces:

Propulsion discipline has

236 of which an SSME

has 25/engine documented

in the formal structured

flight to ground interface

(ICD) system for the

single ground turnaround

facility station (ICD-2-

1A002). Note: The orbiter

element has ten (10) more

facility station ICD’s at

the launch site. Shuttle

System requires the

tracking and managing of

~ 54 different fluids and ~

30 unique fluids are

serviced every flight.

Many of these fluids are

common from one

discipline to another,




vehicle.





realistic and based on the functional

requirements along with assurance the objectives

would be met.

Not considered a need to provide any control as

would have been considered over constraining.

Designers and program managers loss sight of

the objective to build an RLV because of an

overriding focus on meeting the performance

requirements in the absence of any other

structured engineering management processes

used to control such thing as life cycle cost or

any of the other program level one objectives.

Also the STS was optimized at the sub-system

level and not at the overall integrated level.

Electrical functions are custom managed on the

ground and uniquely provided through separate

umbilicals instead of simplifying the flight to

ground interface functions by providing the

electrical management on the vehicle.


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Current Systems


Critical Shortfalls


R 3.6/

R 3.9

25. Easily Supportable

(minimum support

infrastructure required) and

must be compatible with all

LCC requirements by


assumptions

Con’t

The Shuttle has ten (10)

major sub-system













flight. Additionally there

are data bus and


well as unique

instrumentation that have

not been accounted for in

this assessment.


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Appendix VIII

SUMMARY OF

ANNUAL SHUTTLE PAYLOAD DELIVERED OR RETRIEVED FROM SPACE

Table 1. Annual Summary of Space Shuttle Deployed and Retrieved Cargo

Metric Tons

(mt)

Pounds

(lbs.)

Metric

Tons (mt)

Pounds

(lbs.)

Metric

Tons

(mt)

Pounds

(lbs.)

Metric

Tons

(mt)

Pounds

(lbs.)

1981 13.43 29,601 0.00 0 0.00 0 0.00 0

1982 25.03 55,184 6.62 14,585 0.00 0 0.00 0

1983 62.36 137,476 27.19 59,940 1.45 3,192 0.00 0

1984 72.15 159,060 42.76 94,268 0.00 0 1.08 2,381

1985 (2) 110.20 242,956 46.91 103,417 1.01 2,217 0.00 0

1986 35.04 77,258 5.60 12,351 0.00 0 0.00 0

1987 0.00 0 0.00 0 0.00 0 0.00 0

1988 (2) 20.23 44,601 17.02 37,514 0.00 0 0.00 0

1989 (2) 70.47 155,361 59.60 131,397 0.00 0 0.00 0

1990 (2) 45.11 99,450 27.99 61,699 0.00 0 0.00 0

1991 88.37 194,820 56.62 124,820 1.84 4,046 0.00 0

1992 93.62 206,387 27.04 59,613 0.67 1,486 0.00 0

1993 87.39 192,655 30.31 66,826 4.61 10,161 5.25 11,572

1994 72.58 160,009 0.08 171 4.53 9,996 0.00 0

1995 75.40 166,224 21.69 47,812 4.52 9,957 0.53 1,166

1996 64.82 142,898 3.89 8,582 7.71 17,004 6.04 13,321

1997 79.57 175,429 9.49 20,920 4.86 10,724 6.77 14,915

1998 61.01 134,499 15.39 33,931 1.35 2,973 3.09 6,807

1999 38.88 85,704 23.92 52,731 0.00 0 2.52 5,564

2000 64.99 143,274 30.57 67,404 0.00 0 1.30 2,859

2001 81.02 178,619 38.17 84,158 0.00 0 5.51 12,150

2002 58.25 128,410 -- -- -- -- -- --

2003 11.03 24,325 -- -- -- -- -- --

2004 0.00 0 0 0 0.00 0 0.00 0

1330.94 2,934,200 490.86 1,082,139 32.55 71,756 32.09 70,735

Note 2: Does not include Dept of Defense mission payloads flown on the Space Shuttle system from 1985 through 1990.

Total

Note 1: Weights listed are those chargeable to payload; taken from "Space Shuttle Flight Weight Summary," in Space

Shuttle Missions Summary; Book 2, Next 100 Flights; NASA JSC/DA8, Rev. S, May 2002.

Retrieved OnlyTotal Annual

Payload Liftoff Mass (1)

Year

Deployed to SpaceDeployed and

Retrieved


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Appendix IX

Glossary:

List of Acronyms

AHP Analytic Hierarchy Process

APS Auxiliary Propulsion System

APU Auxiliary Power Unit

ASTP Advanced Space Transportation Program

CEC Collaborative Engineering Center

CG Center of Gravity

Crit Criticality

DDT&E Design, Development, Test and Evaluation

Delta V Velocity change

ELV Earth Launch Vehicle

ESMD Exploration Systems Missions Directorate

ET External Tank

ETO Earth-To-Orbit

FRCS Forward Reaction Control System

Gen 3 Third Generation

GH2 Gaseous Hydrogen

GOX Gaseous Oxygen

GRC Glenn Research Center

GSE Ground Support Equipment

HMF Hypergolic Maintenance Facility

H2O Water


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ISS International Space Station

IVHM Integrated Vehicle Health Monitoring

JSC Johnson Space Center

KSC Kennedy Space Center

LCC Life Cycle Cost

LEO Low Earth Orbit

LH2 Liquid Hydrogen

MECO Main Engine(s) Cutoff

MLP Mobile Launch Platform

MPS Main Propulsion System

MSFC Marshall Space Flight Center

NASA National Aeronautics and Space Administration

NH3 Ammonia

NHB NASA Handbook

NPC Non-Propulsive Consumables

NSTS National Space Transportation System

ODS Orbiter Docking System

OMS Orbital Maneuver System

PAD Program Approval Document

PL Payload

POGO launch vehicle induced oscillations (not an acronym; derived from "pogo

stick" analogy)

PPS Power Processing System

PRSD Power Reactants Storage and Distribution

RBCC Rocket Based Combined Cycle


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RCS Reaction Control System

R&D Research and Development

SAIC Science Applications International Corporation

SBS Systems Breakdown Structure

SCAPE Self-Contained Atmospheric Protective Ensemble

SE&I Systems Engineering and Integration

SLWT Super Lightweight Tank

SPST Space Propulsion Synergy Team

SRB Solid Rocket Booster

SRM Solid Rocket Motor

SBS Systems Breakdown Structure

SSME Space Shuttle Main Engine

SSTO Single Stage To Orbit

TBCC Turbine Based Combined Cycle

TBD To Be Determined

TPM Technical Performance Measure

TPS Thermal Protection System

TRL Technology Readiness Level

TSTO Two Stages To Orbit

TV Television

TVC Thrust Vector Control

VAB Vertical Assembly Building

VAFB Vandenburg Air Force Base

WBS Work Breakdown Structure